EP3832105A1 - High power epicyclic gearbox and operation thereof - Google Patents
High power epicyclic gearbox and operation thereof Download PDFInfo
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- EP3832105A1 EP3832105A1 EP20206415.0A EP20206415A EP3832105A1 EP 3832105 A1 EP3832105 A1 EP 3832105A1 EP 20206415 A EP20206415 A EP 20206415A EP 3832105 A1 EP3832105 A1 EP 3832105A1
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- gearbox
- fan
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Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16H—GEARING
- F16H1/00—Toothed gearings for conveying rotary motion
- F16H1/28—Toothed gearings for conveying rotary motion with gears having orbital motion
- F16H1/32—Toothed gearings for conveying rotary motion with gears having orbital motion in which the central axis of the gearing lies inside the periphery of an orbital gear
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/02—Aircraft characterised by the type or position of power plants
- B64D27/10—Aircraft characterised by the type or position of power plants of gas-turbine type
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16H—GEARING
- F16H48/00—Differential gearings
- F16H48/06—Differential gearings with gears having orbital motion
- F16H48/10—Differential gearings with gears having orbital motion with orbital spur gears
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16H—GEARING
- F16H57/00—General details of gearing
- F16H57/02—Gearboxes; Mounting gearing therein
- F16H57/025—Support of gearboxes, e.g. torque arms, or attachment to other devices
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16H—GEARING
- F16H57/00—General details of gearing
- F16H57/08—General details of gearing of gearings with members having orbital motion
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/40—Transmission of power
- F05D2260/403—Transmission of power through the shape of the drive components
- F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
- F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
- F16H—GEARING
- F16H1/00—Toothed gearings for conveying rotary motion
- F16H1/28—Toothed gearings for conveying rotary motion with gears having orbital motion
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the present disclosure relates to gas turbine engines, specifically gas turbine engines for aircraft. Aspects of the present disclosure also relate to an aircraft comprising the gas turbine engine, and a method of operating the gas turbine engine.
- Gas turbine engines for aircraft propulsion have many design factors that affect the overall efficiency and power output or thrust.
- a general aim for a gas turbine engine is to provide low specific fuel consumption (SFC).
- SFC specific fuel consumption
- a larger diameter fan may be used.
- a gearbox is provided having an output to a fan shaft via which the fan is driven.
- the gearbox receives drive from a core shaft connected to a turbine system of the engine core.
- the gearbox allows the fan to operate at a reduced rotational speed compared to if a direct drive were used.
- a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a radial bending stiffness to moment of inertia ratio of: the radial bending stiffness of at least one of the fan shaft
- the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 2.5x10 -2 Nkg -1 m -3 to 6.0 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg -1 m -3 to 3.0 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg -1 m -3 to 0.6 Nkg -1 m -3
- a ratio of the radial bending stiffness of the fan shaft at the output of the gearbox to the moment of inertia of the fan may be greater than or equal to 2.5x10 -2 Nkg -1 m -3 , greater than or equal to 0.05 Nkg -1 m -3 , in the range from 2.5x10 -2 Nkg -1 m -3 to 6.0 Nkg -1 m -3 , in the range from 0.05 Nkg -1 m -3 to 3.0 Nkg -1 m -3 , or in the range from 0.05 Nkg -1 m -3 to 0.6 Nkg -1 m -3 .
- a ratio of the radial bending stiffness of the gearbox support to the moment of inertia of the fan may be greater than or equal to 3.0x10 -2 Nkg -1 m -3 , greater than or equal to 0.06 Nkg -1 m -3 , in the range from 3.0x10 -2 Nkg -1 m -3 to 4.0 Nkg -1 m -3 , in the range from 0.06 Nkg -1 m -3 to 2.0 Nkg -1 m -3 , or in the range from 0.06 Nkg -1 m -3 to 0.48 Nkg -1 m -3 .
- the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 4.00x10 6 N/m.
- the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 3.7x10 7 N/m.
- the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 4.00x10 6 N/m to 1.50x10 9 N/m.
- the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 3.7x10 7 N/m to 1.0x10 9 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 2.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 3.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 1.0x10 7 N/m to 4.0x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 2.0x10 7 N/m to 3.0x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 3.0x10 7 N/m to 2.0x10 8 N/m.
- the diameter of the fan may be in the range from 240 cm to 280 cm.
- the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg -1 m -3 or in the range from 0.05 Nkg -1 m -3 to 4.0 Nkg -1 m -3 .
- the diameter of the fan may be in the range from 330 cm to 380 cm.
- the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.025 Nkg -1 m -3 or in the range from 0.025 Nkg -1 m -3 to 2.0 Nkg -1 m -3 .
- a product e.g. a radial bending stiffness moment of inertia product of: the radial bending stiffiness of at least one of the fan shaft at the ouput of the gearbox and the gearbox support ⁇ the moment of inertia of the fan may be greater than or equal to 2.0x10 14 Nkgm, greater than or equal to 4.0x10 14 Nkgm, greater than or equal to 2.0x10 15 Nkgm, in the range from 2.0x10 14 Nkgm to 1.4x10 18 Nkgm, in the range from 4.0x10 14 Nkgm to 7.0x10 17 Nkgm, or in the range from 2.0x10 15 Nkgm to 7.0x10 17 Nkgm.
- a tilt stiffness to moment of inertia ratio of: the tilt stiffness of at least one of the fan shaft at the output of the gearbox and the gearbox support the moment of inertia of the fan may be greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 2.7x10 -1 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 0.1 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 1.5x10 -2 Nrad -1 kg -1 m -1 .
- a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a tilt stiffness to moment of inertia ratio of: the tilt stiffness of at least one of the fan shaft at the output of the gearbox and
- the tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 2.7x10 -1 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 0.1 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 1.5x10 -2 Nrad -1 kg -1 m -1 .
- a ratio of the tilt stiffness of the fan shaft at the output of the gearbox to the moment of inertia of the fan may be greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 , greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 , in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 2.7x10 -1 Nrad -1 kg -1 m -1 , in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 0.1 Nrad -1 kg -1 m -1 , or in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 1.5x10 -2 Nrad -1 kg -1 m -1 .
- a ratio of the tilt stiffness of the gearbox support to the moment of inertia of the fan may be greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 , greater than or equal to 2.0x10 -3 Nrad -1 kg -1 m -1 , in the range from 1.0x10 -3 Nrad - 1 kg -1 m -1 to 7.0x10 -2 Nrad -1 kg -1 m -1 , in the range from 2.0x10 -3 Nrad -1 kg -1 m -1 to 3.0x10 -2 Nrad -1 kg -1 m -1 , or in the range from 2.0x10 -3 Nrad -1 kg -1 m -1 to 7.0x10 -3 Nrad -1 kg -1 m -1 .
- the tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 7.00x10 4 Nm/rad.
- the tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 9.5x10 5 Nm/rad.
- the tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 7.00x10 4 Nm/rad to 7.00x10 7 Nm/rad.
- the tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 9.5x10 5 Nm/rad to 3.5x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 1.2x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 2.4x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 3.9x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 1.2x10 5 Nm/rad to 2.1x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 2.4x10 5 Nm/rad to 1.6x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 3.9x10 5 Nm/rad to 9.0x10 6 Nm/rad.
- a radial bending stiffness to moment of inertia ratio of: the radial bending stiffness of at least one of the fan shaft at the output of the gearbox and the gearbox support the moment of inertia of the fan may be greater than or equal to 2.5x10 -2 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 2.5x10 -2 Nkg -1 m -3 to 6.0 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg -1 m -3 to 3.0 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg -1 m -3 to 0.6 Nkg -1 m -3 .
- the diameter of the fan may be in the range from 240 cm to 28 cm.
- the tilt stiffness to moment of inertia ratio greater may be than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 or in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 1.45x10 -2 Nrad -1 kg -1 m -1 .
- the diameter of the fan may be in the range from 330 cm to 380 cm.
- the tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 or in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 3.0x10 -2 Nrad -1 kg -1 m -1 .
- a product e.g. the tilt stiffness moment of inertia product of: the tilt stiffness of at least one of the fan shaft at the output of the gearbox and the gearbox support ⁇ the moment of inertia of the fan may be greater than or equal to 3.0x10 12 Nm 3 rad -1 kg, greater than or equal to 6.0x10 12 Nm 3 rad -1 kg, greater than or equal to 2.5x10 13 Nm 3 rad -1 kg, in the range from 3.0x10 12 Nm 3 rad -1 kg to 6.0x10 16 Nm 3 rad -1 kg, in the range from 6.0x10 12 Nm 3 rad -1 kg to 3.0x10 16 Nm 3 rad -1 kg, or in the range from 2.5x10 13 Nm 3 rad -1 kg to 3.0x10 16 Nm 3 rad -1 kg.
- the moment of inertia of the fan may be greater than or equal to 8.3x10 7 kgm 2 .
- the moment of inertia of the fan may be in the range from 7.40x10 7 kgm 2 to 9.00x10 8 kgm 2 .
- the moment of inertia of the fan may be in the range from 8.3x10 7 kgm 2 to 6.5x10 8 kgm 2 .
- the fan blades may be formed from a metallic material.
- the fan blades may be formed from titanium or an aluminium-lithium alloy with a titanium leading edge.
- the fan blades may be formed at least partly from an organic matrix composite.
- the fan blades may have a carbon composite body with a metallic leading edge.
- the fan shaft may connect the output of the gearbox to the fan.
- a gearbox output position may be defined as the point of connection between the fan shaft and the gearbox.
- the fan may have an axial centreline.
- a fan-gearbox axial distance may be defined as the axial distance between the axial position of the gearbox output position and the fan axial centreline.
- the fan-gearbox axial distance multiplied by the moment of inertia of the fan may be greater than or equal to 1.9x10 7 kgm 3 , greater than or equal to 2.9x10 7 kgm 3 , in the range from 1.9x10 7 kgm 3 to 6.2x10 8 kgm 3 , or in the range from 2.9x10 7 kgm 3 to 3.9x10 8 kgm 3 .
- the fan shaft may be defined as the torque transfer component extending from the output of the gearbox to the input to the fan.
- the fan shaft may comprise at least part of a gearbox output shaft and at least part of a fan input shaft.
- the gearbox may be in a star configuration and the output of the gearbox may be a gearbox output position defined as the point of connection between the ring gear and the fan shaft.
- the gearbox may be in a planetary configuration and the output of the gearbox may be a gearbox output position at the interface between the fan shaft and the planet carrier.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; a gearbox support arranged to at least partially support the gearbox within the propulsor, and wherein: the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a radial bending stiffness to moment of inertia ratio of: the radial bending stiffness of at least one of the fan shaft at the output of the gearbox and the gearbox support the moment of inertia of the fan is greater than or equal to 2.5x10 -2 Nkg -1 m -3 .
- the propulsor of the third aspect may have some or all of the features described above with respect to the gas turbine engine of the first aspect, and may be a gas turbine engine in some embodiments.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a gearbox support arranged to at least partially support the gearbox within the propulsor, and wherein: the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a tilt stiffness to moment of inertia ratio of: the tilt stiffness of at least one of the fan shaft at the output of the gearbox and the gearbox support the moment of inertia of the fan is greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 .
- the propulsor of the fourth aspect may have some or all of the features described above with respect to the gas turbine engine of the second aspect, and may be a gas turbine engine in some embodiments.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a gearbox support arranged to at least partially support the gearbox within the propulsor, and wherein the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and wherein:
- the propulsor of the fifth aspect may have some or all of the features described above with respect to the gas turbine engine of the first or second aspect, and may be a gas turbine engine in some embodiments.
- a method of operating a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a radial bending stiffness to moment of inertia ratio of: the radial bending
- the method of the sixth aspect may be a method of operating the gas turbine engine or the propulsor of the first aspect or third aspect respectively. Any of the features, ratios and parameters introduced above in connection with the first or third aspect may also therefore apply to the sixth aspect
- a seventh aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a tilt stiffness to moment of inertia ratio of: the tilt stiffness of at least one of the fan
- the method of the seventh aspect may be a method of operating the gas turbine engine or the propulsor of the second aspect or fourth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the second aspect or fourth aspect may also therefore apply to the seventh aspect.
- a method of operating a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 , and wherein: a radial bending stiffness to moment of inerti
- the method of the eighth aspect may be a method of operating the gas turbine engine or the propulsor of the first, second or fifth aspect. Any of the features, ratios and parameters introduced above in connection with the first, second or fifth aspect may also therefore apply to the eighth aspect.
- the inventor has discovered that either the radial bending/tilt stiffness of one or both of the fan shaft at the gearbox output and the gearbox support need to be low enough for a given moment of inertia of the fan to isolate the gearbox from the transmission of loads from the fan. Rotation of the fan will cause a gyroscopic effect meaning that the fan shaft will tend to maintain a steady direction of its axis of rotation. During manoeuvring of the aircraft to which the gas turbine engine is mounted the orientation of the axis of rotation of the fan shaft will however change. The inventor has found that gyroscopic effects will result in loads being transmitted to the gearbox.
- the inventor has discovered that by using a fan shaft and/or gearbox support having a radial bending/tilt stiffness so that the radial bending/stilt stiffness to moment of inertia ratio defined above is within the specified range the problem of load from the fan being transmitted to the gearbox can be addressed.
- the inventor has found that if the radial bending/tilt stiffness of the fan shaft or the gearbox support shaft is increased so that the ratio is outside of the defined range there is inadequate isolation of the gearbox.
- the inventor has also found that if the fan shaft/gearbox support stiffness were to be decreased so that the ratio is outside of the range above excessive vibration at low modal frequencies would result.
- value ranges for the product of the components of the radial bending stiffness to moment of inertia ratio and the tilt stiffness to moment of inertia ratio may be specified instead of, or as well as, value ranges for the ratios.
- the first aspect introduced above may be formulated as an aspect providing a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a product of: the radial bending stiffness of at least one of the fan shaft at the output of the
- the second aspect introduced above may be formulated as an aspect providing a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x10 7 kgm 2 ; and a product of: the tilt stiffness of at least one of the fan shaft at the output of the gearbox and the
- a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 ; and a first gearbox support strength ratio of: the torsional strength of the gearbox support the radial bending stiffness of the gearbox support ⁇ the cross sectional
- the first gearbox support strength ratio may be greater than or equal to 1.0x10 -2 .
- the first gearbox support strength ratio may be greater than or equal to 7.0x10 -3 .
- the first gearbox support strength ratio may be greater than or equal to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.5x10 -1 .
- the first gearbox support strength ratio may be in the range from 1.0x10 -2 to 1.0x10 -1 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 2.0x10 -2 to 2.5x10 -1 .
- the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 2.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 3.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 1.0x10 7 N/m to 4.0x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 2.0x10 7 N/m to 3.0x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 3.0x10 7 N/m to 2.0x10 8 N/m.
- the torsional strength of the gearbox support may be greater than or equal to 1.60x10 5 Nm.
- the torsional strength of the gearbox support may be greater than or equal to 1.8x10 5 Nm.
- the torsional strength of the gearbox support may be in the range from 1.60x10 5 Nm to 2.00x10 7 Nm.
- the torsional strength of the gearbox support may be in the range from 1.8x10 5 Nm to 1.5x10 6 Nm.
- a planet gear spacing angle in radians may be defined as 2 ⁇ / N , where N is the number of planet gears.
- a second gearbox support strength ratio of: the torsional strength of the gearbox support the tilt stiffness of the gearbox support ⁇ the planet gear spacing angle may be greater than or equal to 1.0x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 1.5x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 1.0x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 3.5.
- the second gearbox support strength ratio may be in the range from 1.5x10 -1 to 1.7.
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 2.5x10 -1 to 3.5.
- a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: a planet gear spacing angle in radians is defined as 2 ⁇ / N , where N is the number of planet gears; the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 ; and a second gearbox support strength ratio
- the second gearbox support strength ratio may be greater than or equal to 1.5x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 3.5.
- the second gearbox support strength ratio may be in the range from 1.5x10 -1 to 1.7.
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 2.5x10 -1 to 3.5.
- the tilt stiffness of the gearbox support may be greater than or equal to 1.2x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 2.4x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 3.9x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 1.2x10 5 Nm/rad to 2.1x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 2.4x10 5 Nm/rad to 1.6x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 3.9x10 5 Nm/rad to 9.0x10 6 Nm/rad.
- the torsional strength of the gearbox support may be greater than or equal to 1.60x10 5 Nm.
- the torsional strength of the gearbox support may be greater than or equal to 1.8x10 5 Nm.
- the torsional strength of the gearbox support may be in the range from 1.60x10 5 Nm to 2.00x10 7 Nm.
- the torsional strength of the gearbox support may be in the range from 1.8x10 5 Nm to 1.5x10 6 Nm.
- a first gearbox support strength ratio of: the torsional strength of the gearbox support the radial bending stiffness of the gearbox support ⁇ the cross sectional area of the gearbox may be greater than or equal to 7.0x10 -3 .
- the first gearbox support strength ratio may be greater than or equal to 1.0x10 -2 .
- the first gearbox support strength ratio may be greater than or equal to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.5x10 -1 .
- the first gearbox support strength ratio may be in the range from 1.0x10 -2 to 1.0x10 -1 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 2.0x10 -2 to 2.5x10 -1 .
- the gearbox may be in a star configuration.
- the planet gear spacing angle may be greater than or equal to 9.0x10 -1 rad.
- the planet gear spacing angle may be in the range between 9.0x10 -1 rad to 2.1 rad.
- the cross sectional area of the gearbox may be greater than or equal to 2.6x10 -1 m 2 .
- the cross sectional area of the gearbox may be in the range from 2.4x10 -1 m 2 to 1.10 m 2 .
- the cross sectional area of the gearbox may be in the range from 2.6x10 -1 m 2 to 9.0x10 -1 m 2 .
- the cross sectional area (CSA) of the gearbox may be defined as the area of the pitch circle of the ring gear.
- a torsional shear stress of the gearbox support, at maximum take-off conditions may be less than or equal to 4.90x10 8 N/m 2 , less than or equal to 3.5x10 8 N/m 2 , in the range from 1.40x10 8 N/m 2 to 4.90x10 8 N/m 2 , or in the range from 2.0x10 8 N/m 2 to 3.5x10 8 N/m 2 .
- a first gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maxiumum take off conditions the radial bending stiffness of the gearbox support may be less than or equal to 4.9x10 -1 m -1 , less than or equal to 20 m -1 , in the range from 3.5x10 -1 m -1 to 4.9x10 1 m -1 , or in the range from 0.70 m -1 to 20 m -1 .
- a second gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maxiumum take off conditions the tilt stiffness of the gearbox support may be less than or equal to 4.1x10 3 rad/m 3 , less than or equal to 1.4x10 3 rad/m 3 , in the range from 6.6 rad/m 3 to 4.1x10 3 rad/m 3 , or in the range from 1.25x10 1 rad/m 3 to 1.4x10 3 rad/m 3 .
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 ; a first gearbox support strength ratio of: the torsional strength of the gearbox support the radial bending stiffness of the gearbox support ⁇ the cross sectional area of the gear
- the propulsor of the eleventh aspect may have some or all of the features described above with respect to the gas turbine engine of the ninth aspect, and may be a gas turbine engine in some embodiments.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 ; and a planet gear spacing angle in radians is defined as 2 ⁇ / N , where N is the number of planet gears; and a second gearbox support strength ratio of
- the propulsor of the twelfth aspect may have some or all of the features described above with respect to the gas turbine engine of the tenth aspect, and may be a gas turbine engine in some embodiments.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 , and wherein: a first gearbox support strength ratio of: the torsional strength of the gearbox support the radial
- the propulsor of the thirteenth aspect may have some or all of the features described above with respect to the gas turbine engine of the ninth aspect and tenth aspects, and may be a gas turbine engine in some embodiments.
- a fourteenth aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 ; and a first gearbox support strength ratio of: the torsional strength of the gearbox support the radial bending stiffness
- the method of the fourteenth aspect may be a method of operating the gas turbine engine or the propulsor of the ninth aspect or eleventh aspect respectively. Any of the features, ratios and parameters introduced above in connection with the ninth aspect or eleventh aspect may also therefore apply to the fourteenth aspect.
- a method of operating a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 ; and a planet gear spacing angle in radians is defined as 2 ⁇ / N , where N is the number of planet gears
- the method of the fifteenth aspect may be a method of operating the gas turbine engine or the propulsor of the tenth aspect or twelfth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the tenth aspect or twelfth aspect may also therefore apply to the fifteenth aspect.
- a sixteenth aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10 -1 m 2 , and
- the method of the sixteenth aspect may be a method of operating the gas turbine engine or the propulsor of the ninth, tenth or thirteenth aspect. Any of the features, ratios and parameters introduced above in connection with the ninth, tenth or thirteenth aspect may also therefore apply to the sixteenth aspect.
- the inventor has discovered that by designing the gearbox support so that the ratio of its torsional strength to its radial bending stiffness and cross sectional area (i.e. the first gearbox support strength ratio) is within the specified range sufficient strength of the support is provided so that the engine is reliable with sufficient stiffness provided to minimise misalignment of the gears in the gearbox and avoid vibration.
- the inventor has found that a similar consideration applies to the ratio of the torsional strength to the tilt stiffness and planet gear spacing angle (i.e. the second gearbox support strength ratio).
- a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: a first gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maximum take off conditions the radial bending stiffness of the gearbox support is less than or equal to 4.9x10 1 m -1 .
- the first gearbox support shear stress ratio may be less than or equal to 20 m -1 .
- the first gearbox support shear stress ratio may be in the range from 3.5x10 -1 m -1 to 4.9x10 1 m -1 .
- the first gearbox support shear stress ratio may be in the range from 0.70 m -1 to 20 m -1 .
- the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 2.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 3.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 1.0x10 7 N/m to 4.0x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 2.0x10 7 N/m to 3x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 3.0x10 7 N/m to 2.0x10 8 N/m.
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be less than or equal to 4.90x10 8 N/m 2 .
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be less than or equal to 3.5x10 8 N/m 2 .
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be in the range from 1.40x10 8 N/m 2 to 4.90x10 8 N/m 2 .
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be in the range from 2.0x10 8 N/m 2 to 3.5x10 8 N/m 2 .
- the diameter of the fan may be in the range from 240 cm to 280 cm.
- the first gearbox support shear stress ratio may be less than or equal to 35 m -1 , or in the range from 0.70 m -1 to 35 m -1 .
- the diameter of the fan may be in the range from 330 cm to 380 cm.
- the first gearbox support shear stress ratio may be less than or equal to 12 m -1 , or in the range from 0.50 m -1 to 12 m -1 .
- a second gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maxiumum take off conditions the tilt stiffness of the gearbox support may be less than or equal to 4.1x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be less than or equal to 1.4x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be in the range from 6.6 rad/m 3 to 4.1x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be in the range from 1.25x10 1 rad/m 3 to 1.4x10 3 rad/m 3 .
- a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox arranged to at least partially support the gearbox within the engine, and: a second gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maximum take off conditions tilt stiffness of the gearbox support is less than or equal to 4.1x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be less than or equal to 1.4x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be in the range from 6.6 rad/m 3 to 4.1x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be in the range from 1.25x10 1 rad/m 3 to 1.4x10 3 rad/m 3 .
- the tilt stiffness of the gearbox support may be greater than or equal to 1.2x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 2.4x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 3.9x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 1.2x10 5 Nm/rad to 2.1x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 2.4x10 5 Nm/rad to 1.6x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 3.9x10 5 Nm/rad to 9.0x10 6 Nm/rad.
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be less than or equal to 4.90x10 8 N/m 2 .
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be less than or equal to 3.5x10 8 N/m 2 .
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be in the range from 1.40x10 8 N/m 2 to 4.90x10 8 N/m 2 .
- the torsional shear stress of the gearbox support, at maximum take-off conditions may be in the range from 2.0x10 8 N/m 2 to 3.5x10 8 N/m 2 .
- the diameter of the fan may be in the range from 240 cm to 280 cm.
- the second gearbox support shear stress ratio may be less than or equal to 2.9x10 3 rad/m 3 , or in the range from 2.9x10 1 rad/m 3 to 2.9x10 3 rad/m 3 .
- the diameter of the fan may be in the range from 330 cm to 380 cm.
- the second gearbox support shear stress ratio may be less than or equal to 7.0x10 2 rad/m 3 or in the range from 1.0x10 1 rad/m 3 to 7.0x10 2 rad/m 3 .
- a first gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maxiumum take off conditions radial bending stiffness of the gearbox support may be less than or equal to 4.9x10 1 m -1 .
- the first gearbox support shear stress ratio may be less than or equal to 20 m -1 .
- the first gearbox support shear stress ratio may be in the range from 3.5x10 -1 m -1 to 4.9x10 1 m -1 .
- the first gearbox support shear stress ratio may be in the range from 0.70 m -1 to 20 m -1 .
- the torsional strength of the gearbox support may be greater than or equal to 1.60x10 5 Nm.
- the torsional strength of the gearbox support may be greater than or equal to 1.8x10 5 Nm.
- the torsional strength of the gearbox support may be in the range from 1.60x10 5 Nm to 2.00x10 7 Nm.
- the torsional strength of the gearbox support may be in the range from 1.8x10 5 Nm to 1.5x10 6 Nm.
- the cross sectional area of the gearbox may be greater than or equal to 2.4x10 -1 m 2 .
- the cross sectional area of the gearbox may be greater than or equal to 2.6x10 -1 m 2 .
- the cross sectional area of the gearbox may be in the range from 2.4x10 -1 m 2 to 1.10 m 2 .
- the cross sectional area of the gearbox may be in the range from 2.6x10 -1 m 2 to 9.0x10 -1 m 2 .
- a planet gear spacing angle in radians may be defined as 2 ⁇ / N , where N is the number of planet gears.
- the planet gear spacing angle may be greater than or equal to 9.0x10 -1 rad.
- the planet gear spacing angle may be in the range between 9.0x10 -1 rad to 2.1 rad.
- a first gearbox support strength ratio of: the torsional strength of the gearbox support the radial bending stiffness of the gearbox support ⁇ the cross sectional area of the gearbox may be greater than or equal to 7.0x10 -3 .
- the first gearbox support strength ratio may be greater than or equal to 1.0x10 -2 .
- the first gearbox support strength ratio may be greater than or equal to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.5x10 -1 .
- the first gearbox support strength ratio may be in the range from 1.0x10 -2 to 1.0x10 -1 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 2.0x10 -2 to 2.5x10 -1 .
- a second gearbox support strength ratio of: the torsional strength of the gearbox support the tilt stiffness of the gearbox support ⁇ the planet gear spacing angle may be greater than or equal to 1.0x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 1.5x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 3.5.
- the second gearbox support strength ratio may be in the range from 1.5x10 -1 to 1.7.
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 2.5x10 -1 to 3.5.
- the torque transmitted through the gearbox support, at maximum take-off conditions may be greater than or equal to 6.00x10 4 Nm, greater than or equal to 7.2x10 4 Nm, in the range from 6.00x10 4 Nm to 5.00x10 5 Nm, or in the range from 7.2x10 4 Nm to 4.2x10 5 Nm.
- the gearbox may be in a planetary configuration.
- the torque transmitted through the gearbox support, at maximum take-off conditions may be greater than or equal to 6.00x10 4 Nm, greater than or equal to 7.2x10 4 Nm, in the range from 6.00x10 4 Nm to 3.00x10 5 Nm, or in the range from 7.2x10 4 Nm to 2.6x10 5 Nm.
- the gearbox may be in a star configuration.
- the torque transmitted through the gearbox support, at maximum take-off conditions may be greater than or equal to 1.10x10 5 Nm, greater than or equal to 1.3x10 5 Nm, in the range from 1.10x10 5 Nm to 5.00x10 5 Nm, or in the range from 1.3x10 5 Nm to 4.2x10 5 Nm.
- the maximum take-off conditions may be as defined anywhere herein, for example may be defined as operating at maximum take-off thrust for the engine at ISA sea level pressure and temperature +15°C with a fan inlet velocity of between 0.25 and 0.27 Mn, and optionally at 0.25 Mn.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: a first gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maximum take off conditions the radial bending stiffness of the gearbox support is less than or equal to 4.9x10 1 m -1 .
- the propulsor of the nineteenth aspect may have some or all of the features described above with respect to the gas turbine engine of the seventeenth aspect, and may be a gas turbine engine in some embodiments.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: a second gearbox support shear stress ratio of: the torsional shear stress of the gearbox support at maximum take off conditions the tilt stiffness of the gearbox support is less than or equal to 4.1x10 3 rad/m 3 .
- the propulsor of the twentieth aspect may have some or all of the features described above with respect to the gas turbine engine of the eighteenth aspect, and may be a gas turbine engine in some embodiments.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein:
- the propulsor of the twenty-first aspect may have some or all of the features described above with respect to the gas turbine engine of the seventeenth aspect and eighteenth aspects, and may be a gas turbine engine in some embodiments.
- a method of operating a gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that, at maximum take-off conditions, a first gearbox support shear stress ratio of: the torsional shear stress of the gearbox support radial bending stiffness of the gearbox support is less than or equal to 4.9x
- the method of the twenty-second aspect may be a method of operating the gas turbine engine or the propulsor of the seventeenth aspect or the nineteenth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the seventeenth aspect or nineteenth aspect may also therefore apply to the twenty-second aspect.
- a method of operating a gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that, at maximum take-off conditions, a second gearbox support shear stress ratio of: the torsional shear stress of the gearbox support the tilt stiffness of the gearbox support is less than or equal to 4.1x10 3
- the method of the twenty-third aspect may be a method of operating the gas turbine engine or the propulsor of the eighteenth aspect or the twentieth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the eighteenth aspect or twentieth aspect may also therefore apply to the twenty-second aspect.
- the twenty-second and twenty-third aspects may be combined.
- a method of operating a gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that, at maximum take-off conditions:
- the method of the previous aspect may be a method of operating the gas turbine engine or the propulsor of the seventeenth, eighteenth or twenty-first aspect. Any of the features, ratios and parameters introduced above in connection with the seventeenth, eighteenth or twenty-first aspect may also therefore apply to this aspect.
- the inventor has discovered that by designing the gearbox support so that the ratio of its torsional shear stress at max take-off to its stiffness (radial bending or tilt stiffness) is within the specified range sufficient margin of strength of the support is provided at the highest load point in the engine operating cycle while the stiffness is suitable to ensure a suitable load share factor.
- a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein: a flight cycle ratio of: the torsional shear stress of the gearbox support at maximum take off MTO conditions the torsional shear stress of the gearbox support at cruise conditions is less than or equal to 3.20.
- the flight cycle ratio may be less than or equal to 2.95.
- the flight cycle ratio may be less than or equal to 2.9.
- the flight cycle ratio may be less than or equal to 2.90.
- the flight cycle ratio may be less than or equal to 2.85.
- the flight cycle ratio may be less than or equal to 2.75.
- the flight cycle ratio may be in the range from 2.10 to 3.20.
- the flight cycle ratio may be in the range from 2.3 to 2.9.
- the torsional shear stress of the gearbox support at maximum take-off conditions may be less than or equal to 4.90x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at maximum take-off conditions may be less than or equal to 2.0x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 1.40x10 8 N/m 2 to 4.9x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 2.0x10 8 N/m 2 to 3.5x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 7.00x10 7 N/m 2 .
- the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 8.2x10 7 N/m 2 .
- the torsional shear stress of the gearbox support at cruise conditions may be in the range from 7.00x10 7 N/m 2 to 1.90x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at cruise conditions may be in the range from 8.2x10 7 N/m 2 to 1.5x10 8 N/m 2 .
- the diameter of the fan may be in the range from 240 cm to 280 cm.
- the flight cycle ratio may be less than or equal to 3.2 or in the range from 2.3 to 3.2.
- the diameter of the fan may be in the range from 330 cm to 380 cm.
- the flight cycle ratio may be less than or equal to 2.8 or in the range from 2.1 to 2.8.
- a product of: the torsional shear stress of the gearbox support at maximum take off conditions ⁇ the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 1.00x10 16 (N/m 2 ) 2 , greater than or equal to 2.05x10 16 (N/m 2 ) 2 , in the range from 1.00x10 16 (N/m 2 ) 2 to 7.50x10 16 (N/m 2 ) 2 , or in the range from 2.05x10 16 (N/m 2 ) 2 to 4.9x10 16 (N/m 2 ) 2 .
- the gearbox may, for example, be in a star configuration.
- a first torque transmission ratio of: the torque transmitted through the gearbox at maximum take off conditions the torque transmitted through the gearbox at cruise conditions may be less than or equal to 3.2.
- the first torque transmission ratio may be less than or equal to 2.95.
- the first torque transmission ratio may be less than or equal to 2.9.
- the first torque transmission ratio may be less than or equal to 2.90.
- the first torque transmission ratio may be less than or equal to 2.85.
- the first torque transmission ratio may be less than or equal to 2.75.
- the first torque transmission ratio may be in the range from 2.1 to 3.2.
- the first torque transmission ratio may be in the range from 2.3 to 2.9.
- a second torque transmission ratio of: the torque transmitted through the gearbox support at maximum take off conditions the torque transmitted through the gearbox support at cruise conditions may be less than or equal to 3.2.
- the second torque transmission ratio may be less than or equal to 2.95.
- the second torque transmission ratio may be less than or equal to 2.9.
- the second torque transmission ratio may be less than or equal to 2.90.
- the second torque transmission ratio may be less than or equal to 2.85.
- the second torque transmission ratio may be less than or equal to 2.75.
- the second torque transmission ratio may be in the range from 2.1 to 3.2.
- the second torque transmission ratio may be in the range from 2.3 to 2.9.
- the torsional strength of the gearbox support may be greater than or equal to 1.60x10 5 Nm.
- the torsional strength of the gearbox support may be greater than or equal to 1.8x10 5 Nm.
- the torsional strength of the gearbox support may be in the range from 1.60x10 5 Nm to 2.00x10 7 Nm.
- the torsional strength of the gearbox support may be in the range from 1.8x10 5 Nm to 1.5x10 6 Nm.
- the gearbox may have a cross sectional area, and a first gearbox support strength ratio of: the torsional strength of the gearbox support the radial bending stiffness of the gearbox support ⁇ the cross sectional area of the gearbox may be greater than or equal to 7.0x10 -3 .
- the first gearbox support strength ratio may be greater than or equal to 1.0x10 -2 .
- the first gearbox support strength ratio may be greater than or equal to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.5x10 -1 .
- the first gearbox support strength ratio may be in the range from 1.0x10 -2 to 1.0x10 -1 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 2.0x10 -2 to 2.5x10 -1 .
- a planet gear spacing angle in radians may be defined as 2 ⁇ / N , where N is the number of planet gears.
- a second gearbox support strength ratio of: the torsional strength of the gearbox support the tilt stiffness of the gearbox support ⁇ the planet gear spacing angle may be greater than or equal to 1.0x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 1.5x10 -1 .
- the second gearbox support strength ratio may be greater than or equal to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 3.5.
- the second gearbox support strength ratio may be in the range from 1.5x10 -1 to 1.7.
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 2.5x10 -1 to 3.5.
- the cross sectional area of the gearbox may be greater than or equal to 2.4x10 -1 m 2 .
- the cross sectional area of the gearbox may be greater than or equal to 2.6x10 -1 m 2 .
- the cross sectional area of the gearbox may be in the range from 2.4x10 -1 m 2 to 1.10 m 2 .
- the cross sectional area of the gearbox may be in the range from 2.6x10 -1 m 2 to 9.0x10 -1 m 2 .
- the planet gear spacing angle may be greater than or equal to 9.0x10 -1 rad.
- the planet gear spacing angle may be in the range between 9.0x10 -1 rad to 2.1 rad.
- the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 2.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 3.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 1.0x10 7 N/m to 4.0x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 2.0x10 7 N/m to 3x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 3.0x10 7 N/m to 2.0x10 8 N/m.
- the tilt stiffness of the gearbox support may be greater than or equal to 1.2x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 2.4x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 3.9x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 1.2x10 5 Nm/rad to 2.1x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 2.4x10 5 Nm/rad to 1.6x10 7 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 3.9x10 5 Nm/rad to 9.0x10 6 Nm/rad.
- the gearbox may be in a star configuration.
- the maximum take-off conditions may be as defined anywhere herein, for example as operating at maximum take-off thrust for the engine at ISA sea level pressure and temperature +15°C with a fan inlet velocity of between 0.25 and 0.27 Mn, and optionally at 0.25 Mn.
- the cruise conditions may be as defined anywhere herein.
- the cruise conditions may mean the conditions at mid-cruise of an aircraft to which the engine is attached.
- the cruise conditions may be conditions experienced by the aircraft and engine at the midpoint between top of climb and start of decent.
- the forward speed of the gas turbine engine at the cruise conditions may be in the range of from Mn 0.75 to Mn 0.85.
- the forward speed of the gas turbine engine at the cruise conditions may be Mn 0.8.
- the cruise conditions may correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11582 m and a forward Mach Number of 0.8.
- the cruise conditions may correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10668 m and a forward Mach Number of 0.85.
- the cruise conditions may correspond to atmospheric conditions at an altitude that is in the range of from 10500 m to 11600 m, and optionally at an altitude of 11000 m.
- a propulsor for an aircraft comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: a flight cycle ratio of: the torsional shear stress of the gearbox support at maximum take off conditions the torsional shear stress of the gearbox support at cruise conditions is less than or equal to 3.20.
- the propulsor of the twenty-fifth aspect may have some or all of the features described above with respect to the gas turbine engine of the twenty-fourth aspect, and may be a gas turbine engine in some embodiments.
- an aircraft comprising the gas turbine engine or the propulsor of the twenty-fourth aspect or the twenty-fifth aspect mounted thereon, wherein the aircraft has a maximum take-off operating condition and a cruise condition.
- a method of operating a gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that a flight cycle ratio of: the torsional shear stress of the gearbox support at maximum takeoff conditions the torsional shear stress of the gearbox support at cruise conditions is less than or equal to 3.20.
- the method of the twenty-sixth aspect may be a method of operating the gas turbine engine or the propulsor of the twenty-fourth aspect or twenty-fifth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the twenty-fourth aspect or twenty-fifth aspect may also therefore apply to the twenty-sixth aspect.
- the method of the twenty-fifth aspect may comprise operating the gas turbine engine to provide propulsion for the aircraft to which it is mounted under maximum take-off conditions.
- the method may further comprise operating the gas turbine engine to provide propulsion during cruise conditions. Cruise conditions and max-take off conditions are as defined elsewhere herein.
- gearbox failure modes including tooth root bending, tooth surface scuffing, tooth surface macro and micro pitting and gearbox vibration - and the extent to which those failure modes have an impact at maximum take-off (MTO) and cruise
- MTO maximum take-off
- the inventor has discovered that providing torsional shear stresses of the gearbox support structure within the specified range provides a low weight gearbox capable of successfully transmitting the largest loads at MTO for limited durations while providing sufficient integrity against other failure mechanisms at cruise for extended durations.
- value ranges for the product of the components of the flight cycle ratio may be specified instead of, or as well as, value ranges for the ratio.
- the twenty-fourth aspect introduced above may be formulated as an aspect providing a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein: a product (e.g.
- a flight cycle product of: the torsional shear stress of the gearbox support at maximum take off conditions ⁇ the torsional shear stress of the gearbox support at cruise conditions is greater than or equal to 1.00x10 16 (N/m 2 ) 2 , greater than or equal to 2.05x10 16 (N/m 2 ) 2 , in the range from 1.00x10 16 (N/m 2 ) 2 to 7.50x10 16 (N/m 2 ) 2 , or in the range from 2.05x10 16 (N/m 2 ) 2 to 4.9x10 16 (N/m 2 ) 2 .
- the gearbox may be in a star configuration.
- the turbine may be a first turbine
- the compressor may be a first compressor
- the core shaft may be a first core shaft.
- the engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor.
- the second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft
- the gearbox may have a gear ratio in any range disclosed herein, for example a gear ratio in the range from 3.2 to 4.5, and optionally from 3.2 to 4.0.
- the gas turbine engine may have a specific thrust in any range disclosed herein, for example a specific thrust in the range from 70 to 90 NKg -1 .
- the gas turbine engine may have a bypass ratio at cruise conditions in any range disclosed herein, for example in the range from 12.5 to 18, and optionally from 13 to 16.
- the fan may have a fan diameter greater than 240 cm and less than or equal to 380 cm.
- the fan may have a fan diameter greater than 300 cm and less than or equal to 380 cm.
- the diameter of the fan may be in the range from 240 cm to 280 cm.
- the diameter of the fan may be in the range from 330 cm to 380 cm.
- the cross sectional area (CSA) of the gearbox may be defined as the area of the pitch circle of the ring gear.
- the method of any of the aspects defined above may further comprise driving the gearbox with an input torque of:
- a limit on the values that X can take that is expressed as "X is greater than or equal to Y” can alternatively be expressed as "1/X is less than or equal to 1/Y". Any of the ratios or parameters defined in the aspects and statements above may therefore be expressed as "1/X is less than or equal to 1/Y" rather than "X is greater than or equal to Y". In such cases, zero can be taken as a lower bound.
- the inventor has found that a relatively low stiffness may be provided to isolate the gearbox from damaging loads being transmitted into it.
- the inventor has found that if the stiffness of the gearbox support is decreased outside of the ranges specified herein there are resulting problems with dynamic effects such as lateral vibrations - in particular, a stiffness less that is within ranges defined herein has been found to allow vibrations at low modal frequencies to be reduced or avoided (the skilled person would appreciate that the lower modal vibrations have larger amplitudes/deflections than the higher modes, and so are more important to avoid). This may be a function of the size of the gearbox and its configuration.
- the inventor has also found that the maximum stiffnesses provided by the ranges defined herein allows the reduction or avoidance of damaging loads being transmitted to the gearbox from the fan. This may similarly vary according to size and gearbox configuration.
- the inventor has found that decreasing the radial bending and/or tilt stiffness of the fan shaft (at the input to the fan or the output of the gearbox) outside of the ranges defined herein would lead to undesirable dynamic effects such as lateral vibration.
- the minimum stiffness defined by the ranged specified herein allows vibrations at low modal frequencies to be reduced or avoided (the skilled person would appreciate that the lower modal vibrations have larger amplitudes/deflections than the higher modes, and so are more important to avoid). This may be a function of the size of the gearbox and its configuration.
- an upper limit of the fan shaft radial bending and/or tilt stiffness is affected by the fundamental properties of the material or materials from which it is made.
- a maximum stiffness is affected by the engineering limit of the material from which it is made.
- the materials from which the fan shaft is made may, for example, have a Young's modulus in the range from 100 to 250 GPa, or 105 to 215 GPa, and optionally around 210 GPa - different grades of steel, or other types of metal, may be selected to achieve different stiffnesses for the same size and geometry.
- steels with a Young's modulus in the range 190 to 215 GPa, titanium alloys with a Young's modulus in the range 105 to 120 GPa, or a metal such as titanium with a Young's modulus of around 110 GPa may be used in various embodiments.
- the inventor has found increasing the stiffness outside of the ranges defined herein using materials such as these would lead to excessive weight with no practical gain in performance.
- the inventor has discovered that the range of the torsional strength of the gearbox support defined herein provides a desired level of torque capacity to provide sufficient reliability, while not being so great as to add excessive weight to the engine with no practical performance gain.
- Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor.
- a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.
- the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.
- the input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear.
- the core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).
- the output from the gearbox may be directly to a fan shaft, or indirectly to the fan shaft, for example via a spur shaft and/or gear.
- the gas turbine engine as described and/or claimed herein may have any suitable general architecture.
- the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts.
- the turbine connected to the core shaft may be a first turbine
- the compressor connected to the core shaft may be a first compressor
- the core shaft may be a first core shaft.
- the engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor.
- the second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft
- the second compressor may be positioned axially downstream of the first compressor.
- the second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.
- the gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above).
- the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above).
- the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.
- the gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used.
- the gearbox may be a "planetary” or “star” gearbox, as described in more detail elsewhere herein.
- the gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2.
- the gear ratio may be, for example, between any two of the values in the previous sentence.
- the gearbox may be a "star” gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.
- a combustor may be provided axially downstream of the fan and compressor(s).
- the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided.
- the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided.
- the combustor may be provided upstream of the turbine(s).
- each compressor may comprise any number of stages, for example multiple stages.
- Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable).
- the row of rotor blades and the row of stator vanes may be axially offset from each other.
- each turbine may comprise any number of stages, for example multiple stages.
- Each stage may comprise a row of rotor blades and a row of stator vanes.
- the row of rotor blades and the row of stator vanes may be axially offset from each other.
- Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position.
- the ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25.
- the ratio of the radius of the fanblade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio.
- the radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade.
- the hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.
- the radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge.
- the fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches).
- the fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.
- the rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm.
- the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.
- the fan In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity U tip .
- the work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow.
- a fan tip loading may be defined as dH/U tip 2 , where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and U tip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed).
- the fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4.
- the fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.
- Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions.
- the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20.
- the bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14.
- the bypass duct may be substantially annular.
- the bypass duct may be radially outside the core engine.
- the radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.
- the overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor).
- the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75.
- the overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.
- Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s or 80 Nkg -1 s.
- the specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg -1 s to 100 Nkg -1 s, or 85 Nkg -1 s to 95 Nkg -1 s.
- Such engines may be particularly efficient in comparison with conventional gas turbine engines.
- a gas turbine engine as described and/or claimed herein may have any desired maximum thrust.
- a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN.
- the maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).
- a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN.
- the thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.
- the temperature of the flow at the entry to the high pressure turbine may be particularly high.
- This temperature which may be referred to as TET
- TET may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane.
- the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K.
- the TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).
- the maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K.
- the maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K.
- the maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.
- MTO maximum take-off
- a fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials.
- at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre.
- at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material.
- the fan blade may comprise at least two regions manufactured using different materials.
- the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade.
- a leading edge may, for example, be manufactured using titanium or a titanium-based alloy.
- the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.
- a fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction.
- the fan blades may be attached to the central portion in any desired manner.
- each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc).
- a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc.
- the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring.
- any suitable method may be used to manufacture such a bladed disc or bladed ring.
- at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.
- variable area nozzle may allow the exit area of the bypass duct to be varied in use.
- the general principles of the present disclosure may apply to engines with or without a VAN.
- the fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.
- a maximum take-off (MTO) condition has the conventional meaning.
- Maximum take-off conditions may be defined as operating the engine at International Standard Atmosphere (ISA) sea level pressure and temperature conditions +15°C at maximum take-off thrust at end of runway, which is typically defined at an aircraft speed of around 0.25Mn, or between around 0.24 and 0.27 Mn.
- Maximum take-off conditions for the engine may therefore be defined as operating the engine at a maximum take-off thrust (for example maximum throttle) for the engine at ISA sea level pressure and temperature +15°C with a fan inlet velocity of 0.25 Mn.
- cruise conditions have the conventional meaning and would be readily understood by the skilled person.
- the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached.
- mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint - in terms of time and/or distance- between top of climb and start of descent.
- Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i.e.
- cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude).
- the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.
- the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.
- the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m.
- the cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.
- the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m).
- the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m).
- a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein.
- cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.
- an aircraft comprising a gas turbine engine as described and/or claimed herein.
- the aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.
- a method of operating a gas turbine engine as described and/or claimed herein may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).
- a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein.
- the operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.
- gearbox configurations disclosed herein Whilst in the arrangements described herein the source of drive for the propulsive fan is provided by a gas turbine engine, the skilled person will appreciate the applicability of the gearbox configurations disclosed herein to other forms of aircraft propulsor comprising alternative drive types.
- the above-mentioned gearbox arrangements may be utilised in aircraft propulsors comprising a propulsive fan driven by an electric motor.
- the electric motor may be configured to operate at higher rotational speeds and thus may have a lower rotor diameter and may be more power-dense.
- the gearbox configurations of the aforesaid aspects may be employed to reduce the rotational input speed for the fan or propeller to allow it to operate in a more favourable efficiency regime.
- an electric propulsion unit for an aircraft comprising an electric machine configured to drive a propulsive fan via a gearbox, the gearbox and/or its inputs/outputs/supports and/or the structure by which the fan shaft driving the fan is supported being as described and/or claimed herein.
- the term “axial plane” denotes a plane extending along the length of an engine, parallel to and containing an axial centreline of the engine
- the term “radial plane” denotes a plane extending perpendicular to the axial centreline of the engine, so including all radial lines at the axial position of the radial plane.
- Axial planes may also be referred to as longitudinal planes, as they extend along the length of the engine. A radial distance or an axial distance is therefore a distance in a radial or axial plane, respectively.
- FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9.
- the engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B.
- the gas turbine engine 10 comprises a core 11 that receives the core airflow A.
- the engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20.
- a nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18.
- the bypass airflow B flows through the bypass duct 22.
- the fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.
- the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place.
- the compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted.
- the resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust.
- the high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27.
- the fan 23 generally provides the majority of the propulsive thrust.
- the epicyclic gearbox 30 is a reduction gearbox.
- FIG. 2 An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2 .
- the low pressure turbine 19 (see Figure 1 ) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30.
- a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30 Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34.
- the planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis.
- the planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9.
- an annulus or ring gear 38 Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.
- the linkages 36 may be referred to as a fan shaft 36, the fan shaft 36 optionally comprising two or more shaft portions coupled together.
- the fan shaft 36 may comprise a gearbox output shaft portion 36a extending from the gearbox 30 and a fan portion 36b extending between the gearbox output shaft portion and the fan 23.
- the gearbox 30 is a planetary gearbox and the gearbox output shaft portion 36a is connected to the planet carrier 34 - it may therefore be referred to as a carrier output shaft 36a.
- the gearbox output shaft portion 36a may be connected to the ring gear 38 - it may therefore be referred to as a ring output shaft 36a.
- the fan portion 36b of the fan shaft 36 connects the gearbox output shaft portion 36a to the fan 23.
- the output of the gearbox 30 is therefore transferred to the fan 23, to rotate the fan, via the fan shaft 36.
- the fan shaft 36 may comprise a single component, or more than two components. Unless otherwise indicated or apparent to the skilled person, anything described with respect to an engine 10 with a star gearbox 30 may equally be applied to an engine with a planetary gearbox 30, and vice versa.
- low pressure turbine and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23).
- the "low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the "intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.
- the epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3 .
- Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in Figure 3 .
- Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.
- the epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed.
- the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38.
- the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.
- any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10.
- the connections (such as the linkages 36, 40 in the Figure 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility as defined or claimed elsewhere herein.
- any suitable arrangement of the bearings between rotating and stationary parts of the engine may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2 .
- the gearbox 30 has a star arrangement (described above)
- the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2 (for example as described in connection with other embodiments disclosed herein which have a star gearbox arrangement).
- the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.
- gearbox styles for example star or planetary
- support structures for example star or planetary
- input and output shaft arrangement for example star or planetary
- bearing locations for example star or planetary
- the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).
- additional and/or alternative components e.g. the intermediate pressure compressor and/or a booster compressor.
- gas turbine engines to which the present disclosure may be applied may have alternative configurations.
- such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts.
- the gas turbine engine shown in Figure 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core engine nozzle 20.
- this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle.
- One or both nozzles may have a fixed or variable area.
- the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.
- the geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in Figure 1 ), and a circumferential direction (perpendicular to the page in the Figure 1 view).
- the axial, radial and circumferential directions are mutually perpendicular.
- stiffnesses may be used herein:
- a radial bending stiffness is a measure of deformation for a given force applied in any one selected radial direction (i.e. any direction perpendicular to and passing through the engine axis).
- the radial bending stiffness is defined with reference to Figure 4 in terms of the deformation of a cantilevered beam 401.
- a force, F applied at the free end of the beam in a direction perpendicular to the longitudinal axis of the beam causes a linear perpendicular deformation, ⁇ .
- the radial bending stiffness is the force applied for a given linear deformation i.e. F/ ⁇ .
- the radial bending stiffness is taken relative to the rotational axis of the engine 9, and so relates to the resistance to linear deformation in a radial direction of the engine caused by a radial force.
- the beam, or equivalent cantilevered component extends along the axis of rotation of the engine, the force, F, is applied perpendicular to the axis of rotation of the engine, along any radial direction, and the displacement ⁇ is measured perpendicular to the axis of rotation, along the line of action of the force.
- the radial bending stiffness as defined herein has SI units of N/m.
- the radial bending stiffness is taken to be a free-body stiffness i.e. stiffness measured for a component in isolation in a cantilever configuration, without other components present which may affect its stiffness.
- a tilt stiffness is defined with reference to Figure 5 , which shows the resulting deformation of a cantilevered beam 401 under a moment M applied at its free end.
- the tilt stiffness is a measure of the resistance to rotation of a point on the component at which a moment is applied.
- an applied moment at the free end of the cantilevered beam induces a constant curvature along the length of the beam between its free and fixed ends.
- the applied moment M causes a rotation ⁇ of the point at which it is applied.
- the angle ⁇ is constant along the length of the component.
- the tilt stiffness as defined herein therefore has SI units of Nm/rad.
- Torsional stiffness is a measure of deformation for a given torque.
- Figure 6 illustrates the definition of the torsional stiffness of a shaft 401 or other body.
- a torque, ⁇ applied to the free end of the beam causes a rotational deformation, ⁇ (e.g. twist) along the length of the beam.
- the torsional stiffness is the torque applied for a given angle of twist i.e. ⁇ / ⁇ .
- the torsional stiffness has SI units of Nm/rad.
- Torque which may also be referred to as moment, is the rotational equivalent of linear force, and can be thought of as a twist to an object.
- the net torque on a body determines the rate of change of the body's angular momentum.
- Moment of inertia is a quantity that determines the torque needed for a desired angular acceleration of a body about a rotational axis - this is substantially equivalent to how mass determines the force needed for a particular acceleration.
- Moment of inertia depends on the body's mass distribution and the axis chosen, with larger moments requiring more torque to change the body's rotation rate.
- Moment of inertia has dimensions of [mass]x[distance] 2 and may be expressed in units of kilogram meter squared (kg. m 2 ).
- a fan of a gas turbine engine (and more generally a fan rotor of the gas turbine engine comprising the fan disc and blades, and optionally also the fan shaft and/or other related components) is constrained to rotate in only one plane - a plane perpendicular to the engine axis - and that the fan's moment of inertia can therefore be defined by a single, scaler, value.
- the fan's moment of inertia about the engine axis can therefore be measured or defined using any standard methodology
- Shear stress is the component of stress coplanar with a material cross section; the stress tending to produce shear. Shear stress arises from a force vector component parallel to the cross section of the material. Shear stress may be defined as the external force acting on an object or surface parallel to the slope or plane in which it lies.
- a shearing stress When a shaft or other body is subjected to a torque, or twisting, a shearing stress is produced in the body - the shearing stress may be referred to as a torsional shear stress as it results from a torque.
- the shear stress varies from zero along the axis of the torque rotation to a maximum at the part of the body furthest from the axis; a radial distance from the axis at which the shear stress is to be measured is therefore selected.
- a mid-height of the component of which the torsional shear stress is being measured is selected. For example, if the component has an outer radius of 20 cm, a mid-height position of 10 cm from the axis 9 is selected.
- Torsional shear stress has dimensions of [force]/[distance] 2 and may be expressed in units of Pascals (Pa) or Newtons/metre squared (N/m 2 ).
- Shear strength is the strength of a body against the body failing in shear - Shear strength is a material's ability to resist forces that can cause the internal structure of the material to slide against itself.
- a shear load is a force that tends to produce a sliding failure on a material along a plane that is parallel to the direction of the force.
- the ultimate shear strength of a material is the maximum shear stress that may be sustained before the material will rupture.
- the proof strength of a material is the stress at which a particular degree of permanent deformation occurs - for example, a 0.2% proof strength is the stress at which a 0.2% permanent deformation occurs.
- torsional shear strength The strength of the body against shearing when a rotational, or twisting, shear load is applied may be referred to as torsional shear strength.
- shear strength of a component is important for selecting the dimensions and materials to be used for the manufacture or construction of the component. Shear strength has the same units as torsional shear stress, as it is a maximum supportable shear stress; it may therefore be expressed in units of Pa or N/m 2 .
- the listed shear strength of a material is the 0.2% proof shear strength of the material unless otherwise specified. The skilled person would appreciate that this is lower than the ultimate shear strength.
- Shear strength of a material generally varies with temperature.
- the shear strength as used herein may be defined at room temperature.
- shear strength generally decreases gradually with increasing temperature until a threshold temperature is reached, beyond which the material strength decreases rapidly.
- the shear strength may therefore be approximately constant below the threshold temperature.
- the threshold temperature may be around 400-500°C for various steel grades. As the temperatures in and around the gearbox 30 generally do not exceed 120°C, well below the threshold temperature for likely materials for gearbox, gearbox support structure, and shaft construction, the shear strength may not be significantly different from the room temperature shear strength - the choice of a specific temperature for strength assessment may therefore have little effect.
- the torsional strength of a specific component is defined as the ability of a component to withstand an applied torque without failure (i.e. without yield or structural failure).
- the torsional strength is therefore a measure of the torque capacity of a component.
- the torsional strength may be measured by applying a varying torque to a component and determining the torque at which the component fails.
- the torsional strength can be considered as the torque applied to the component at the point where the shear strength of that component is reached causing it to fail.
- the torsional strength as defined herein has dimensions of [force] x [distance] and may be expressed in units of Nm.
- stiffnesses and other parameters relating to embodiments described herein are provided below for ease of understanding.
- Figure 7 shows a region of the engine core 11 around the gearbox in close up.
- the same reference numbers have been used for components corresponding to those shown in Figures 1 to 3 .
- the gearbox 30 has a star arrangement, in which the ring gear 38 is coupled to the fan shaft 36 and the carrier 34 is held in a fixed position relative to the static structure of the engine core (e.g. relative to the stationary supporting structure 24).
- the fan shaft 36 is mounted within the engine by a fan shaft mounting structure 503.
- the fan shaft mounting structure 503 comprises at least two bearings connected to or otherwise in engagement with the fan shaft at points spaced apart axially along the length of the engine.
- the fan shaft mounting structure 503 may take a number of different forms, and may comprise one or more separate supporting structures provided to support the fan shaft. It may also include other structures provided to support the fan shaft such as inter-shaft bearings. It therefore includes any supporting structure that extends between a bearing in contact with the fan shaft and a stationary structure of the engine (e.g. of the engine core).
- the fan shaft mounting structure 503 comprises two bearings, a first supporting bearing 506a and a second supporting bearing 506b, via which it is coupled to the fan shaft 36.
- the supporting bearings 506a, 506a are spaced apart along the axial length of the fan shaft 36.
- both supporting bearings 506a, 506b are provided at positions that are forward of the gearbox 30.
- one of the two supporting bearing 506a, 506b used to support the fan shaft 36 may be located at a position rearward of the gearbox 30.
- more than two supporting bearings may be provided as part of the fan shaft mounting structure or fan shaft mounting structure.
- the engine core 11 comprises a gearbox support 40 (corresponding to the linkage described with reference to Figure 2 ) arranged to support or mount (e.g. to at least partially support or mount) the gearbox 30 in a fixed position within the engine.
- the gearbox support is coupled at a first end to the stationary supporting structure 24 which extends across the core duct 502 carrying the core airflow A as illustrated in Figure 7 .
- the stationary support structure 24 is an engine section stator (ESS) that acts as both a structural component to provide a stationary mounting for core components such as the gearbox support, and as a guide vane provided to direct airflow from the fan 23.
- ESS engine section stator
- the stationary supporting structure 24 may comprise a strut extending across the core gas flow path and a separate stator vane provided to direct airflow.
- the gearbox support 40 is coupled at a second end to the planet carrier 34. The gearbox support 40 therefore acts against rotation of the planet carrier 34 relative to the static structure of the engine core (e.g. relative to the stationary supporting structure 24).
- the gearbox support 40 is coupled to the ring gear 38 so as to resist its rotation relative to the static structure of the engine core (e.g. relative to the stationary supporting structure 24).
- the gearbox support 40 is defined between the point at which it connects to the gearbox (e.g. to the planet carrier in the presently described arrangement) and a point at which it connects to the stationary supporting structure 24.
- the gearbox support may be formed by any number of separate components providing a coupling between those two points.
- the gearbox support 40 has a degree of flexibility characterized by its radial bending stiffness and its tilt stiffness.
- the radial bending stiffness of the gearbox support 40 is defined with reference to Figures 8 and 9 .
- the radial bending stiffness can be considered to represent the resistance of the gearbox support to a force applied to it in a radial direction of the engine.
- the radial bending stiffness is determined by treating the gearbox support 40 as a free body that is fixed at its point of connection with the stationary supporting structure 24 of the engine, and which has a radial force F 1 applied at its point of connection with the gearbox 30 as illustrated in Figure 8 . Deformation of the gearbox support 40 caused by the applied force F 1 is illustrated in Figure 9 , with the shape of the support with no force applied shown in broken lines for comparison.
- the radial bending stiffness is defined in terms of the radial displacement, ⁇ 1 , of the gearbox support at the position at which the force F 1 is applied.
- the force, F 1 is shown radially towards the engine axis 9 in Figure 8 , but could equivalently be a force radially away from the engine axis 9.
- the radial bending stiffness of the gearbox support 40 is therefore given by F 1 / ⁇ 1 .
- the gearbox 30 is shown for reference in Figure 9 remaining in a stationary position despite the deformation of the gearbox support 40 to which it is connected - this is for illustration purposes only and reflects the gearbox support 40 being treated as a free body in order to determine the stiffness.
- the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x10 7 N/m, and optionally greater than or equal to 2.0x10 7 N/m or greater than or equal to 3.0x10 7 N/m.
- the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x10 7 N/m or greater than or equal to 2.7x10 7 N/m. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness of the gearbox support may be greater than or equal to 3.2x10 7 N/m or greater than or equal to 4.0x10 7 N/m
- the radial bending stiffness of the gearbox support may be in the range from 1.0x10 7 N/m to 4.0x10 8 N/m, and optionally in the range from 2.0x10 7 N/m to 3x10 8 N/m or further optionally in the range from 3.0x10 7 N/m to 2.0x10 8 N/m.
- the radial bending stiffness of the gearbox support may be in the range from 1.0x10 7 N/m to 3.1x10 8 N/m and optionally in the range from 2.7x10 7 N/m to 3.7x10 7 N/m (and may be equal to 3.2x10 7 N/m).
- the radial bending stiffness of the gearbox support may be in the range from 3.2x10 7 N/m to 4.0x10 8 N/m, and optionally in the range from 4.0x10 7 N/m to 5.0x10 7 N/m (and may be equal to 4.5x10 7 N/m).
- the tilt stiffness of the gearbox support 40 is defined with reference to Figures 10 and 11 .
- the tilt stiffness can be considered to represent the resistance of the gearbox support to an applied moment.
- the tilt stiffness is determined by treating the gearbox support 40 as a free body that is fixed at its point of connection with the stationary supporting structure 24.
- a moment M 1 is applied at the point of connection between the gearbox support 40 and the gearbox 30 as illustrated in Figure 10 .
- Deformation of the gearbox support 40 caused by the applied moment M 1 is illustrated in Figure 11 , with the shape of the support with no moment applied shown in broken lines.
- the tilt stiffness is defined in terms of the angular displacement, ⁇ 1 , of the gearbox support 40 at the position at which the moment M 1 is applied. The tilt stiffness is therefore given by M 1 / ⁇ 1 .
- the gearbox 30 is again shown for reference in Figure 11 remaining in a stationary position despite the deformation of the gearbox support 40 to which it is connected.
- the tilt stiffness of the gearbox support may be greater than or equal to 1.2x10 5 Nm/rad, and optionally greater than or equal to 2.4x10 5 Nm/rad or optionally greater than or equal to 3.9x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be greater than or equal to 1.2x10 5 Nm/rad or greater than or equal to 4.5x10 5 Nm/rad. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness of the gearbox support may be greater than or equal to 5.0x10 5 Nm/rad or greater than or equal to 6.0x10 5 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 1.2x10 5 Nm/rad to 2.1x10 7 Nm/rad, and optionally in the range from 2.4x10 5 Nm/rad to 1.6x10 7 Nm/rad, and further optionally in the range from 3.9x10 5 Nm/rad to 9.0x10 6 Nm/rad.
- the tilt stiffness of the gearbox support may be in the range from 1.2x10 5 Nm/rad to 7.0x10 6 Nm/rad and optionally in the range from 4.5x10 5 Nm/rad to 6.5x10 5 Nm/rad (and may be equal to 5.5x10 5 Nm/rad).
- the tilt stiffness of the gearbox support may be in the range from 5.0x10 5 Nm/rad to 2.1x10 7 Nm/rad, and optionally in the range from 6.0x10 5 Nm/rad to 2.6x10 6 Nm/rad (and may be equal to 1.6x10 6 Nm/rad).
- the gearbox support is provided to locate the gearbox within the engine. In order to resist relative rotation of the gearbox compared to the stationary structure of the engine a torque is transmitted through the gearbox support. This varies over the operating cycle of the engine as different levels of torque are transmitted through the gearbox (as defined elsewhere herein) at different stages of the flight cycle of the aircraft to which the engine is mounted.
- the torque transmitted through the gearbox support is defined as the torque at the point of connection between the gearbox support 40 and the gearbox 30.
- the torsional strength of the gearbox support 40 is defined as the level of torque applied at the point of connection between the gearbox support 40 and the gearbox 20 that would result in failure of the gearbox support.
- the gearbox support When measuring the torsional strength or torque transmitted through the gearbox support 40 the gearbox support is considered to be a free body that is fixed at its point of connection with the stationary supporting structure 24 of the engine with a torque applied at the point of connection to the gearbox (i.e. in a similar manner as the application of the force or moment used to determine the radial bending stiffness and tilt stiffness of the gearbox support).
- the point of connection between the gearbox support 40 and the gearbox 30 is the point of connection to the planet carrier 34 of the gearbox. Where the gearbox is in a planetary configuration, this point of connection would be between the gearbox support 40 and the ring gear 38.
- the torsional strength of the gearbox support may be greater than or equal to 1.60x10 5 Nm, and optionally greater than or equal to 1.8x10 5 Nm.
- the torsional strength of the gearbox support may be greater than or equal to 1.8x10 5 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torsional strength of the gearbox support may be greater than or equal to 4.0x10 5 Nm or greater than or equal to 5.5x10 5 Nm.
- the torsional strength of the gearbox support may be in the range from 1.60x10 5 Nm to 2.00x10 7 Nm, and optionally in the range from 1.8x10 5 Nm to 1.5x10 6 Nm.
- the torsional strength of the gearbox support may be in the range from 1.8x10 5 Nm to 7.0x10 5 Nm and optionally in the range from 1.8x10 5 Nm to 2.6x10 5 Nm (and may be equal to 2.2x10 5 Nm).
- the torsional strength of the gearbox support may be in the range from 4.0x10 5 Nm to 2.0x10 7 Nm and optionally in the range from 5.5x10 5 Nm to 7.5x10 5 Nm (and may be equal to 6.5x10 5 Nm).
- the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 6.00x10 4 Nm, and optionally greater than or equal to 7.2x10 4 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 7.0x10 4 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.8x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 6.00x10 4 Nm to 5.00x10 5 Nm, and optionally in the range from 7.2x10 4 Nm to 4.2x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 7.0x10 4 Nm to 1.9x10 5 Nm.
- torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.8x10 5 Nm to 4.5x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.10x10 5 Nm, and optionally greater than or equal to 1.3x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.2x10 5 Nm or greater than or equal to 1.4x10 5 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 3.0x10 5 Nm or is greater than or equal to 3.4x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.10x10 5 Nm to 5.00x10 5 Nm, and optionally in the range from 1.3x10 5 Nm to 4.2x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.2x10 5 Nm to 1.9x10 5 Nm and optionally in the range from 1.4x10 5 Nm to 1.8x10 5 Nm (and may be equal to 1.6x10 5 Nm).
- torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 3.0x10 5 Nm to 4.5x10 5 Nm and optionally in the range from 3.4x10 5 Nm to 4.2x10 5 Nm (and may be equal to 3.8x10 5 Nm).
- the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 6.00x10 4 Nm, and optionally greater than or equal to 7.2x10 4 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 7.0x10 4 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.8x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 6.00x10 4 Nm to 3.00x10 5 Nm, and optionally in the range from 7.2x10 4 Nm to 2.6x10 5 Nm.
- the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 7.0x10 4 Nm to 1.1x10 5 Nm.
- torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.8x10 5 Nm to 2.5x10 5 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.00x10 4 Nm, and optionally greater than or equal to 2.2x10 4 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.2x10 4 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 6.8x10 4 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.00x10 4 Nm to 2.00x10 5 Nm, and optionally in the range from 2.2x10 4 Nm to 1.7x10 5 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.2x10 4 Nm to 6.6x10 4 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be in the range from 6.8x10 4 Nm to 1.8x10 5 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 4.00x10 4 Nm, and optionally greater than or equal to 4.8x10 4 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 4.5x10 4 Nm or greater than or equal to 5.0x10 4 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or the gearbox is in a star configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 1.2x10 5 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be in the range from 4.00x10 4 Nm to 2.00x10 5 Nm, and optionally in the range from 4.8x10 4 Nm to 1.7x10 5 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be in the range from 4.5x10 4 Nm to 6.6x10 4 Nm and optionally in the range from 5.0x10 4 Nm to 6.0x10 4 Nm (and may be equal to 5.5x10 4 Nm).
- torque transmitted through the gearbox support at cruise conditions maybe in the range from 1.2x10 5 Nm to 1.8x10 5 Nm and optionally in the range from 1.2x10 5 Nm to 1.8x10 5 Nm (and may be equal to 1.5x10 5 Nm).
- the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.00x10 4 Nm, and optionally greater than or equal to 2.2x10 4 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.2x10 4 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 6.8x10 4 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.00x10 4 Nm to 1.30x10 5 Nm, and optionally in the range from 2.2x10 4 Nm to 1.1x10 5 Nm.
- the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.2x10 4 Nm to 3.3x10 4 Nm.
- torque transmitted through the gearbox support at cruise conditions may be in the range from 6.8x10 4 Nm to 1.1x10 5 Nm.
- the gearbox support 40 has a torsional shear stress that also represents the resistance of the gearbox support 40 to a torque applied by the gearbox 30.
- the torsional shear stress is as defined elsewhere herein.
- the torsional shear stress of the gearbox support at maximum take-off conditions may be less than or equal to 4.90x10 8 N/m 2 , and optionally less than or equal to 3.5x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 1.40x10 8 N/m 2 to 4.90x10 8 N/m 2 , and optionally in the range from 2.0x10 8 N/m 2 to 3.5x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 2.3x10 8 N/m 2 to 3.7x10 8 N/m 2 (and may be equal to 2.5x10 8 N/m 2 ).
- the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 7.00x10 7 N/m 2 , and optionally greater than or equal to 8.2x10 7 N/m 2 .
- the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 8.0x10 7 N/m 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 8.5x10 7 N/m 2 or greater than or equal to 9.0x10 7 N/m 2 .
- the torsional shear stress of the gearbox support at cruise conditions may be in the range from 7.00x10 7 N/m 2 to 1.90x10 8 N/m 2 , and optionally in the range from 8.2x10 7 N/m 2 to 1.5x10 8 N/m 2 .
- the torsional shear stress of the gearbox support at cruise conditions may be in the range from 8.0x10 7 N/m 2 to 1.5x10 8 N/m 2 and optionally in the range from 8.0x10 7 N/m 2 to 9.2x10 7 N/m 2 (and may be equal to 8.6x10 7 N/m 2 ).
- the torsional shear stress of the gearbox support at cruise conditions may be in the range from 8.5x10 7 N/m 2 to 1.9x10 8 N/m 2 and optionally in the range from 9.0x10 7 Nm 2 to 1.2x10 8 N/m 2 (and may be equal to 9.6x10 7 N/m 2 ).
- the maximum take-off conditions and cruise conditions referred to in this section may be as defined elsewhere herein.
- the shear stress may be measured through the highest loaded plane section of the component (i.e. the gearbox support). It excludes the effects of stress concentrations at small radii and holes.
- the effective radius also affects the measurement of the shear as well as the area.
- the fan 23 has a moment of inertial I F .
- the moment of inertia of the fan is measured based on the total mass of the rotor forming the fan, i.e. including the total mass of the plurality of fan blades, the fan hub and any support arm or other linkages provided to connect the fan to the fan shaft.
- the moment of inertial therefore includes all rotating components apart from the fan shaft.
- the moment of inertia is the mass moment of inertia or rotational inertia of the fan with respect to rotation around the principal rotational axis 9 of the engine. Rotation of the fan will cause a gyroscopic effect meaning that the fan shaft will tend to maintain a steady direction of its axis of rotation.
- the moment of inertia of the fan may be greater than or equal to 7.40x10 7 kgm 2 , and optionally greater than or equal to 8.3x10 7 kgm 2 .
- the moment of inertia of the fan may be greater than or equal to 7.4x10 7 kgm 2 or 8.6x10 7 kgm 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the moment of inertia of the fan may be greater than or equal to 3.0x10 8 kgm 2 or 4.0x10 8 kgm 2 .
- the moment of inertia of the fan may be in the range from 7.40x10 7 kgm 2 to 9.00x10 8 kgm 2 , and optionally in the range from 8.3x10 7 kgm 2 to 6.5x10 8 kgm 2 .
- the moment of inertia of the fan may be in the range from 7.4x10 7 kgm 2 to 1.5x10 8 kgm 2 and optionally in the range from 8.6x10 7 kgm 2 to 9.6x10 7 kgm 2 (and may be equal to 9.1x10 7 kgm 2 ).
- the moment of inertia of the fan may be in the range from 3.0x10 8 kgm 2 to 9.0x10 8 kgm 2 and optionally in the range from 4.0x10 8 kgm 2 to 5.0x10 8 kgm 2 (and may be equal to 4.5x10 8 kgm 2 ).
- a fan-gearbox axial distance 110 is defined as the axial distance between the axial position P of the gearbox output position (also labelled X in Figures 15 to 18 ) and the fan axial centreline Q.
- the fan axial centreline is defined as the axial midpoint of the fan blades forming the fan.
- the gearbox output position is defined as the point of connection between the fan shaft 36 and the gearbox. This may be defined differently for different types of gearbox as is described below.
- the fan-gearbox axial distance may be greater than or equal to 0.35 m, and optionally greater than or equal to 0.37 m.
- the fan-gearbox axial distance may be greater than or equal to 0.38 m or greater than or equal to 0.40 m. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan-gearbox axial distance may be greater than or equal to 0.48 m or greater than or equal to 0.50 m.
- the fan-gearbox axial distance may be in the range from 0.35 m to 0.8 m, and optionally in the range from 0.37 m to 0.75 m.
- the fan-gearbox axial distance may be in the range from 0.38 m to 0.65 m and optionally in the range from 0.40 m to 0.44 m (and may be equal to 0.42 m).
- the fan-gearbox axial distance may be in the range from 0.48 m to 0.8 m and optionally in the range from 0.50 m to 0.66 m (and may be equal to 0.58 m).
- the fan shaft 36 is defined as the torque transfer component that extends from the output of the gearbox to the fan input. It therefore includes part or all of any gearbox output shaft and fan input shaft that may be provided between those points. For the purposes of defining the stiffness of the fan shaft 36 it is considered to extend between a fan input position and a gearbox output position, and includes all of the torque transfer components between those points. It does not therefore include any components of the gearbox (e.g. the planet carrier or connecting plate coupled to it) which transmit discrete forces, rather than the fan shaft torque.
- the gearbox output position therefore may be defined as the point of connection between the fan shaft 36 and the gearbox 30.
- the fan input position may be defined as the point of connection between the fan shaft 36 and the fan.
- the gearbox output position is defined as the point of connection 702 between the ring gear 38 and the fan shaft 36. More specifically, it is the point of connection to the annulus of the ring gear 38 (with any connection component extending from the outer surface of the annulus being considered to be part of the ring gear). Where the point of connection is formed by an interface extending in a direction having an axial component, the point of connection is considered to be the axial centreline of that interface as illustrated in Figure 7 .
- the fan shaft 36 includes all torque transmitting components up to the point of connection 702 with the ring gear 38. It therefore includes any flexible portions or linkages 704 of the fan shaft 36 that may be provided, and any connection(s) 706 (e.g. spline connections) between them.
- the gearbox output position is again defined as the point of connection between the fan shaft 36 and the gearbox 30.
- Figure 13 shows a carrier comprising a forward plate 34a and rearward plate 34b, with a plurality of pins 33 extending between them and on which the planet gears are mounted.
- the fan shaft 36 is connected to the forward plate 34a via a spline connection 708.
- the gearbox output position is taken as any point on the interface between the fan shaft 36 and the forward plate 34a.
- the forward plate 34a is considered to transmit discrete forces, rather than a single torque, and so is taken to be part of the gearbox 30 rather than the fan shaft.
- Figure 13 shows only one example of a type of connection between the fan shaft and planet carrier 34.
- the gearbox output position is still taken to be at the interface between components transmitting a torque (i.e. that are part of the fan shaft) and those transmitting discrete forces (e.g. that are part of the gearbox).
- the spline connection 708 is only one example of a connection that may be formed between the fan shaft and gearbox (i.e. between the fan shaft and the forward plate 34b in the presently described arrangement).
- the interface which forms the gearbox output position may be formed by, for example, a curvic connection, a bolted joint or other toothed or mechanically dogged arrangement.
- the fan input position is defined as a point on the fan shaft at the axial midpoint of the interface between the fan and the fan shaft.
- the fan 23 comprises a support arm 23a (as can be seen, for example, in Figure 7 ) arranged to connect the fan 23 to the fan shaft 36.
- the support arm 23a is connected to the fan shaft by a spline coupling 36c that extends along the length of a portion of the fan shaft 36.
- the fan input position is defined as the axial midpoint of the spline coupling.
- the spline coupling shown is only one example of a coupling that may form the interface between the fan and fan shaft.
- a curvic connection for example, a bolted joint or other toothed or mechanically dogged arrangement may be used.
- a flange coupling may be provided between the support arm 23a and the fan shaft 36.
- the support arm can be connected at the rear of the fan hub.
- the fan input position is the axial midpoint of the flange coupling.
- Figure 14 illustrates an arrangement in which an alternative coupling is provided between the fan 23 and the fan shaft 36.
- the fan 23 is coupled to the fan shaft 36 via a support arm 23a.
- a flange coupling 36d is provided between the support arm 23a and the fan shaft 36.
- the support arm 23a is connected at the rear of the fan hub.
- the flange coupling 36d may be a curvic coupling. In other embodiments, other forms of flange coupling may be provided.
- the fan input position is the axial midpoint (labelled Y) of the flange coupling.
- the fan shaft 36 has a degree of flexibility characterized by its radial bending stiffness and tilt stiffness.
- the stiffness of the fan shaft where it couples to the gearbox 30 is defined with reference to Figures 15 to 18 .
- the radial bending stiffness of the fan shaft 36 at the output of the gearbox 30 is measured by applying a force F 2 to the fan shaft at the gearbox output position defined above (illustrated in Figure 15 ).
- the fan shaft 36 is treated as a free body, and is held fixed at the position of all of bearing positions at which it is supported i.e. the first and second bearings 506a, 506b in the arrangement of Figure 15 .
- force F 2 the fan shaft 36 deforms so that the gearbox output position is displaced by a distance of ⁇ 2 (as illustrated in Figure 16 ).
- the radial bending stiffness of the fan shaft 36 at the output of the gearbox is then given by F 2 / ⁇ 2 .
- the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 4.00x10 6 N/m, and optionally greater than or equal to 3.7x10 7 N/m.
- the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 3.7x10 7 N/m. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 3.9x10 7 N/m or greater than or equal to 5.0x10 7 N/m.
- the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 4.00x10 6 N/m to 1.5x10 9 N/m, and optionally in the range from 3.7x10 7 N/m to 1.0x10 9 N/m.
- the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 3.7x10 7 N/m to 5.0x10 8 N/m and optionally in the range from 3.7x10 7 N/m to 4.3x10 7 N/m (and may be equal to 4.0x10 7 N/m).
- the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 3.9x10 7 N/m to 1.5x10 9 N/m, and optionally in the range from 5.0x10 7 N/m to 9.0x10 7 N/m (and may be equal to 7.0x10 7 N/m).
- the tilt stiffness of the fan shaft 36 at the output of the gearbox 30 is measured by applying a moment M 2 to the fan shaft at the gearbox output position defined above (illustrated in Figure 17 ).
- the fan shaft 36 is again treated as a free body, and is held fixed at the position of all of the bearing positions at which it is supported i.e. the first and second bearings 506a, 506b in the arrangement of Figure 15 .
- moment M 2 the fan shaft 36 deforms so that the gearbox output position is displaced by an angular displacement of ⁇ 2 as illustrated in Figure 17 .
- the tilt stiffness of the fan shaft 36 at the output of the gearbox is then given by M 2 / ⁇ 2 .
- the tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 7.00x10 4 Nm/rad, and optionally greater than or equal to 9.5x10 5 Nm/rad.
- the tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 9.5x10 5 , and optionally may be greater than or equal to 9.5x10 5 Nm/rad. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness of the fan shaft output of the gearbox may be greater than or equal to 1.1x10 6 Nm/rad and optionally may be greater than or equal to 2.0x10 6 Nm/rad .
- the tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 7.00x10 4 Nm/rad to 7.00x10 7 Nm/rad, and optionally in the range from 9.5x10 5 Nm/rad to 3.5x10 7 Nm/rad.
- the tilt stiffness of the output of the gearbox may be in the range from 9.5x10 5 Nm/rad to 2.0x10 7 Nm/rad and optionally in the range from 9.5x10 5 Nm/rad to 2.4x10 6 Nm/rad (and may be equal to 1.2x10 6 Nm/rad).
- the tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 1.1x10 6 Nm/rad to 7.0x10 7 Nm/rad, and optionally in the range from 2.0x10 6 Nm/rad to 5.2x10 6 Nm/rad (and may be equal to 3.6x10 6 Nm/rad).
- the gearbox provides a torque conversion between the toque at its input (i.e. the core shaft) and its output (i.e. the fan shaft).
- the torque transmitted through the gearbox is defined as the torque at the output position of the gearbox (the output position being as defined elsewhere herein).
- the torque transmitted through the gearbox varies over the operating cycle of the engine.
- the torque transmitted through the gearbox at maximum take-off conditions may be greater than or equal to 7.00x10 4 Nm, and optionally greater than or equal to 1.0x10 5 Nm.
- the torque transmitted through the gearbox at maximum take-off conditions may be greater than or equal to 1.1x10 5 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox at maximum take-off conditions may be greater than or equal to 1.5x10 5 Nm or greater than or equal to 2.0x10 5 Nm.
- the torque transmitted through the gearbox at maximum take-off conditions may be in the range from 7.00x10 4 Nm to 5.00x10 5 Nm, and optionally in the range from 1.0x10 5 Nm to 3.5x10 5 Nm.
- the torque transmitted through the gearbox at maximum take-off conditions may be in the range from 1.1x10 5 Nm to 1.5x10 5 Nm and optionally in the range from 1.1x10 5 Nm to 1.3x10 5 Nm (and may be equal to 1.2x10 5 Nm).
- the torque transmitted through the gearbox at maximum take-off conditions may be in the range from 1.5x10 5 Nm to 5.0x10 5 Nm and optionally in the range from 2.0x10 5 Nm to 3.8x10 5 Nm (and may be equal to 2.9x10 5 Nm).
- the torque transmitted through the gearbox at cruise conditions may be greater than or equal to 2.30x10 4 Nm, and optionally greater than or equal to 3.1x10 4 Nm.
- the torque transmitted through the gearbox at cruise conditions may be greater than or equal to 3.2x10 4 Nm or greater than or equal to 3.8x10 4 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox at cruise conditions may be greater than or equal to 7.3x10 4 Nm or greater than or equal to 9.8x10 4 Nm.
- the torque transmitted through the gearbox at cruise conditions may be in the range from 2.30x10 4 Nm to 1.80x10 5 Nm, and optionally in the range from 3.1x10 4 Nm to 1.5x10 5 Nm.
- the torque transmitted through the gearbox at cruise conditions may be in the range from 3.2x10 4 Nm to 7.2x10 4 Nm and optionally in the range from 3.8x10 4 Nm to 4.6x10 4 Nm (and may be equal to 4.2x10 4 Nm).
- torque transmitted through the gearbox at cruise conditions may be in the range from 7.3x10 4 Nm to 1.8x10 5 Nm and optionally in the range from 9.8x10 4 Nm to 1.4x10 5 Nm (and may be equal to 1.1x10 5 Nm).
- the cross sectional area (CSA) of the gearbox is defined as the area of the pitch circle of the ring gear.
- the pitch circle of a gear is an imaginary circle that rolls without slipping with the pitch circle of any other gear with which the first gear is meshed.
- the pitch circle passes through the points where the teeth of two gears meet as the meshed gears rotate - the pitch circle of a gear generally passes through a mid-point of the length of the teeth of the gear.
- the CSA of the gearbox can be found by measuring the pitch circle diameter (PCD) of the gear.
- the PCD can be roughly estimated by taking the average of the diameter between tips of the gear teeth and the diameter between bases of the gear teeth.
- the CSA of the gearbox may be greater than or equal to 2.4x10 -1 m 2 and optionally greater than or equal to 2.6x10 -1 m 2 .
- the CSA of the gearbox may be greater than or equal to 2.4x10 -1 m 2 or greater than or equal to 2.5x10 -1 m 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the CSA of the gearbox may be greater than or equal to 4.5x10 -1 m 2 or greater than or equal to 5.5x10 -1 m 2 .
- the CSA of the gearbox may be in the range from 2.4x10 -1 m 2 to 1.10 m 2 , and optionally in the range from 2.6x10 -1 m 2 to 9.0x10 -1 m 2 .
- the CSA of the gearbox may be in the range from 2.4x10 -1 m 2 to 5.0x10 -1 m 2 and optionally in the range from 2.5x10 -1 m 2 to 3.4x10 -1 m 2 (and may be equal to 2.9x10 -1 m 2 ).
- the CSA of the gearbox may be in the range from 4.5x10 -1 m 2 to 1.1 m 2 , and optionally in the range from 5.5x10 -1 m 2 to 6.4x10 -1 m 2 (and may be equal to 5.9x10 -1 m 2 ).
- a planet gear spacing angle in radians ( ⁇ ) is defined as 2 ⁇ / N, where N is the number of planet gears 32 provided in the gearbox.
- the planet gear spacing angle is illustrated in Figure 3 .
- the planet gearbox spacing angle corresponds to the average angle (in rad) between all the adjacent pairs of planet gears.
- the planet gear spacing angle ( ⁇ ) may be greater than or equal to 9.0x10 -1 rad, and optionally in the range between 9.0x10 -1 rad to 2.1 rad, and further optionally in the range between 1.1 rad and 1.3 rad (and may be equal to 1.26 rad)
- a radial bending stiffness to moment of inertia ratio may be defined as: the radial bending stiffness of at least one of the fan shaft 36 at the ouput of the gearbox and the gearbox support 40 the moment of inertia of the fan 23
- the radial bending stiffness to moment of inertia ratio may be greater than or equal to 2.5x10 -2 Nkg -1 m -3 (i.e. (N/m)/(kg.m 2 )) and optionally greater than or equal to 0.05 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg -1 m -3 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.025 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 2.5x10 -2 Nkg -1 m -3 to 6.0 Nkg -1 m -3 , and optionally in the range from 0.05 Nkg -1 m -3 to 3.0 Nkg -1 m -3 , and further optionally in the range from 0.05 Nkg -1 m -3 to 0.6 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg -1 m -3 to 4.0 Nkg -1 m -3 .
- the radial bending stiffness to moment of inertia ratio may be in the range from 0.025 Nkg -1 m -3 to 2.0 Nkg -1 m -3 .
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the radial bending stiffness to moment of inertia product) may be defined as: the radial bending stiffness of at least one of the fan shaft 36 at the ouput of the gearbox and the gearbox support 40 ⁇ the moment of inertia of the fan 23
- the radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x10 14 Nkgm (i.e. (N/m).(kg.m 2 )), and optionally greater than or equal to 4.0x10 14 Nkgm, and further optionally greater than or equal to 2.0x10 15 Nkgm.
- the radial bending stiffness to moment of inertia product may be greater than or equal to 1.5x10 15 Nkgm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x10 15 Nkgm.
- the radial bending stiffness to moment of inertia product may be in the range from 2.0x10 14 Nkgm to 1.4x10 18 Nkgm, and optionally in the range from 4.0x10 14 Nkgm to 7.0x10 17 Nkgm, and further optionally in the range from 2.0x10 15 Nkgm to 7.0x10 17 Nkgm.
- the radial bending stiffness to moment of inertia product may be in the range from 1.5x10 15 Nkgm to 1.3x10 17 Nkgm.
- the radial bending stiffness to moment of inertia product may be in the range from 2.0x10 15 Nkgm to 1.4x10 18 Nkgm.
- a fan shaft radial bending stiffness to moment of inertia ratio may be defined as: the radial bending stiffness of the fan shaft 36 at the ouput of the gearbox the moment of inertia of the fan 23
- the fan shaft radial bending stiffness to moment of inertia ratio may be greater than or equal to 2.5x10 -2 Nkg -1 m -3 (i.e. (N/m)/(kg.m 2 )), and optionally greater than or equal to 0.05 (N/m)/(kg.m 2 ).
- the fan shaft radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg -1 m -3 or greater than or equal to 0.4 Nkg -1 m -3 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.025 Nkg -1 m -3 or optionally greater than or equal to 0.06 Nkg -1 m -3 .
- the fan shaft radial bending stiffness to moment of inertia ratio may be in the range from 2.5x10 -2 Nkg -1 m -3 to 6.0 Nkg -1 m -3 , and optionally in the range from 0.05 Nkg -1 m -3 to 3.0 Nkg -1 m -3 , and further optionally in the range from 0.05 Nkg -1 m -3 to 0.6 Nkg -1 m -3 .
- the fan shaft radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg -1 m -3 to 0.6 Nkg -1 m -3 and optionally in the range from 0.4 Nkg -1 m -3 to 0.5 Nkg -1 m -3 (and may be equal to 0.44 Nkg -1 m -3 ).
- the fan shaft radial bending stiffness to moment of inertia ratio may be in the range from 0.025 Nkg -1 m -3 to 0.6 Nkg -1 m -3 , and optionally in the range from 0.06 Nkg -1 m -3 to 0.26 Nkg -1 m -3 (and may be equal to 0.16 Nkg -1 m -3 ).
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the fan shaft radial bending stiffness to moment of inertia product) may be defined as: the radial bending stiffness of the fan shaft 36 at the ouput of the gearbox ⁇ the moment of inertia of the fan 23
- the fan shaft radial bending stiffness to moment of inertia product may be greater than or equal to 3.0x10 14 Nkgm (i.e. (N/m).(kg.m 2 )), and optionally greater than or equal to 6.0x10 14 Nkgm, and further optionally greater than or equal to 2.0x10 15 Nkgm.
- the fan shaft radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x10 15 Nkgm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft radial bending stiffness to moment of inertia product may be greater than or equal to 2.3x10 15 Nkgm.
- the fan shaft radial bending stiffness to moment of inertia product may be in the range from 3.0x10 14 Nkgm to 1.4x10 18 Nkgm, and optionally in the range from 6.0x10 14 Nkgm to 7.0x10 17 Nkgm, and further optionally in the range from 2.0x10 15 Nkgm to 70x10 17 Nkgm.
- the fan shaft radial bending stiffness to moment of inertia product may be in the range from 2.0x10 15 Nkgm to 7.5x10 16 Nkgm.
- the fan shaft radial bending stiffness to moment of inertia product may be in the range from 2.3x10 15 Nkgm to 1.4x10 18 Nkgm.
- a gearbox support radial bending stiffness to moment of inertia ratio may be defined as the radial bending stiffness of the gearbox support 40 the moment of inertia of the fan 23
- the gearbox support radial bending stiffness to moment of inertia ratio may be greater than or equal to 3.0x10 -2 Nkg -1 m -3 (i.e. (N/m)/(kg.m 2 )), and optionally greater than or equal to 0.06 Nkg -1 m -3 .
- the gearbox support radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.06 Nkg -1 m -3 or greater than or equal to 0.25 Nkg -1 m -3 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.03 Nkg -1 m -3 or greater than or equal to 0.06 Nkg -1 m -3 .
- the gearbox support radial bending stiffness to moment of inertia ratio may be in the range from 3.0x10 -2 Nkg -1 m -3 to 4.0 Nkg -1 m -3 , and optionally in the range from 0.06 Nkg -1 m -3 to 2.0 Nkg -1 m -3 , and further optionally in the range from 0.06 Nkg -1 m -3 to 0.48 Nkg -1 m -3 .
- the gearbox support radial bending stiffness to moment of inertia ratio may be in the range from 0.06 Nkg -1 m -3 to 4.0 Nkg -1 m -3 and optionally in the range from 0.25 Nkg -1 m -3 to 0.45 Nkg -1 m -3 (and may be equal to 0.35 Nkg -1 m -3 ).
- the gearbox support radial bending stiffness to moment of inertia ratio may be in the range from 0.03 Nkg -1 m -3 to 2.0 Nkg -1 m -3 , and optionally in the range from 0.06 Nkg -1 m -3 to 0.6 Nkg -1 m -3 (and may be equal to 0.1) Nkg -1 m -3 .
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the gearbox support radial bending stiffness to moment of inertia product) may be defined as: the radial bending stiffness of the gearbox support (40) ⁇ the moment of inertia of the fan (23)
- the gearbox support radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x10 14 Nkgm (i.e. (N/m).(kg.m 2 )), and optionally greater than or equal to 4.0x10 14 Nkgm or further optionally greater than or equal to 2.0x10 15 Nkgm.
- the gearbox support radial bending stiffness to moment of inertia product may be greater than or equal to 1.5x10 15 Nkgm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x10 15 Nkgm.
- the gearbox support radial bending stiffness to moment of inertia product may be in the range from 2.0x10 14 Nkgm to 3.0x10 17 Nkgm, and optionally in the range from 4.0x10 14 Nkgm to 1.3x10 17 Nkgm, and yet further optionally in the range from 2.0x10 15 Nkgm to 1.3x10 17 Nkgm.
- the gearbox support radial bending stiffness to moment of inertia product may be in the range from 1.5x10 15 Nkgm to 1.3x10 17 Nkgm.
- the gearbox support radial bending stiffness to moment of inertia product may be in the range from 2.0x10 15 Nkgm to 3.0x10 17 Nkgm.
- a tilt stiffness to moment of inertia ratio may be defined as: the tilt stiffness of at least one of the fan shaft 36 at the output of the gearbox and the gearbox support 40 the moment of inertia of the fan 23
- the tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 (i.e. (Nm/rad)/(kg.m 2 )), and optionally greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10 -1 Nrad -1 kg -1 m -1 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 2.7x10 -1 Nrad -1 kg -1 m -1 , and optionally in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 0.1 Nrad -1 kg -1 m -1 , and further optionally in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 1.5x10 -2 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -1 Nrad -1 kg -1 m -1 to 1.45x10 -2 Nrad -1 kg -1 m -1 .
- the tilt stiffness to moment of inertia ratio may be in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 3.0x10 -2 Nrad -1 kg -1 m -1 .
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the tilt stiffness to moment of inertia product) may be defined as: the tilt stiffness of at least one of the fan shaft 36 at the ouput of the gearbox and the gearbox support 40 ⁇ the moment of inertia of the fan 23
- the tilt stiffness to moment of inertia product may be greater than or equal to 3.0x10 12 Nm 3 rad -1 kg (i.e. (Nm/rad).(kg.m 2 )), and optionally greater than or equal to 6.0x10 12 Nm 3 rad -1 kg, and further optionally greater than or equal to 2.5x10 13 Nm 3 rad -1 kg.
- the tilt stiffness to moment of inertia product may be greater than or equal to 2.0x10 13 Nm 3 rad -1 kg. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness to moment of inertia product may be greater than or equal to 2.5x10 13 Nm 3 rad -1 kg.
- the tilt stiffness to moment of inertia product may be in the range from 3.0x10 12 Nm 3 rad -1 kg to 6.0x10 16 Nm 3 rad -1 kg, and optionally in the range from 6.0x10 12 Nm 3 rad -1 kg to 3.0x10 16 Nm 3 rad -1 kg, and further optionally in the range from 2.5x10 13 Nm 3 rad -1 kg to 3.0x10 16 Nm 3 rad -1 kg.
- the tilt stiffness to moment of inertia product may be in the range from 2.0x10 13 Nm 3 rad -1 kg to 4.0x10 15 Nm 3 rad -1 kg.
- the tilt stiffness to moment of inertia product may be in the range from 2.5x10 13 Nm 3 rad -1 kg to 6.0x10 16 Nm 3 rad -1 kg.
- a fan shaft tilt stiffness to moment of inertia ratio may be defined as: the tilt stiffness of the fan shaft 36 at the output of the gearbox the moment of inertia of the fan 23
- the fan shaft tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 (i.e. (Nm/rad)/(kg.m 2 )), and optionally greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 .
- the fan shaft tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 or greater than or equal to 3.0x10 -3 Nrad -1 kg -1 m -1 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10 -4 Nrad -1 kg -1 m -1 or greater than or equal to 0.7x10 -3 Nrad -1 kg -1 m -1 .
- the fan shaft tilt stiffness to moment of inertia ratio may be in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 0.27 Nrad -1 kg -1 m -1 , and optionally in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 0.1 Nrad -1 kg -1 m -1 , and further optionally in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 1.5x10 -2 Nrad -1 kg -1 m -1 .
- the fan shaft tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 1.45x10 -2 Nrad -1 kg -1 m -1 and optionally in the range from 3.0x10 -3 Nrad -1 kg -1 m -1 to 2.3x10 -2 Nrad -1 kg -1 m -1 (and may be equal to 1.3x10 -2 Nrad -1 kg -1 m -1 ).
- the fan shaft tilt stiffness to moment of inertia ratio may be in the range from 4.0x10 -4 Nrad -1 kg -1 m -1 to 1.4x10 -2 Nrad -1 kg -1 m -1 , and optionally in the range from 7.0x10 -3 Nrad -1 kg -1 m -1 to 9.0x10 -3 Nrad -1 kg -1 m -1 (and may be equal to 8.0x10 -3 Nrad -1 kg -1 m -1 ).
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the fan shaft tilt stiffness to moment of inertia product) may be defined as: the tilt stiffness of the fan shart 36 at the ouput of the gearbox ⁇ the moment of inertia of the fan 23
- the fan shaft tilt stiffness to moment of inertia product may be greater than or equal to 5.0x10 12 Nm 3 rad -1 kg (i.e. (Nm/rad).(kg.m 2 )), and optionally greater than or equal to 1.0x10 13 Nm 3 rad -1 kg or greater than or equal to 6.0x10 13 Nm 3 rad -1 kg.
- the fan shaft tilt stiffness to moment of inertia product may be greater than or equal to 6.0x10 13 Nm 3 rad -1 kg.
- the fan shaft tilt stiffness to moment of inertia product may be greater than or equal to 1.0x10 14 Nm 3 rad -1 kg.
- the fan shaft tilt stiffness to moment of inertia product may be in the range from 5.0x10 12 Nm 3 rad -1 kg to 6.0x10 16 Nm 3 rad -1 kg, and optionally in the range from 1.0x10 13 Nm 3 rad -1 kg to 3.0x10 16 Nm 3 rad -1 kg, and further optionally in the range from 6.0x10 13 Nm 3 rad -1 kg to 3.0x10 16 Nm 3 rad -1 kg.
- the fan shaft tilt stiffness to moment of inertia product may be in the range from 6.0x10 13 Nm 3 rad -1 kg to 3.0x10 15 Nm 3 rad -1 kg.
- the fan shaft tilt stiffness to moment of inertia product may be in the range from 1.0x10 14 Nm 3 rad -1 kg to 6.0x10 16 Nm 3 rad -1 kg.
- a gearbox support tilt stiffness to moment of inertia ratio may be defined as: the tilt stiffness of the gearbox support 40 the moment of inertia of the fan 23
- the gearbox support tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 (i.e. (Nm/rad)/(kg.m 2 )), and optionally greater than or equal to 2.0x10 -3 Nrad -1 kg -1 m -1 .
- the gearbox support tilt stiffness to moment of inertia ratio may be greater than or equal to 2.0x10 -3 Nrad -l1 kg -1 m -1 or greater than or equal to 5.0x10 -3 Nrad -1 kg -1 m -1 .
- the gearbox support tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10 -3 Nrad -1 kg -1 m -1 or greater than or equal to 2.6x10 -3 Nrad -1 kg -1 m -1 .
- the gearbox support tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 7.0x10 -2 Nrad -1 kg -1 m -1 . and optionally in the range from 2.0x10 -3 Nrad -1 kg -1 m -1 to 3.0x10 -2 Nrad -1 kg -1 m -1 , and further optionally in the range from 2.0x10 -3 Nrad -1 kg -1 m -1 to 7.0x10 -3 Nrad -1 kg -1 m -1 .
- the gearbox support tilt stiffness to moment of inertia ratio may be in the range from 2.0x10 -3 Nrad -1 kg -1 m -1 to 7.2x10 -3 Nrad -1 kg -1 m -1 and optionally in the range from 5.0x10 -3 Nrad -1 kg -1 m -1 to 7.0x10 -3 Nrad -1 kg -1 m -1 (and may be equal to 6.0x10 -3 Nrad -1 kg -1 m -1 ).
- the gearbox support tilt stiffness to moment of inertia ratio may be in the range from 1.0x10 -3 Nrad -1 kg -1 m -1 to 3.0x10 -2 Nrad -1 kg -1 m -1 , and optionally in the range from 2.6x10 -3 Nrad -1 kg -1 m -1 to 4.6x10 -3 Nrad -1 kg -1 m -1 (and may be equal to 3.6x10 -3 Nrad -1 kg -1 m -1 ).
- gearbox support tilt stiffness to moment of inertia ratio a product of the parameters making up that ratio may be defined.
- This product (referred to as the gearbox support tilt stiffness to moment of inertia product) may be defined as: the tilt stiffness of gearbox support 40 ⁇ the moment of inertia of the fan 23
- the gearbox support tilt stiffness to moment of inertia product may be greater than or equal to 3.0x10 12 Nm 3 rad -1 kg (i.e. (Nm/rad).(kg.m 2 )), and optionally greater than or equal to 6.0x10 12 Nm 3 rad -1 kg, and further optionally greater than or equal to 2.5x10 13 Nm 3 rad -1 kg.
- the gearbox support tilt stiffness to moment of inertia product may be greater than or equal to 2.0x10 13 Nm 3 rad -1 kg. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support tilt stiffness to moment of inertia product may be greater than or equal to 2.5x10 13 Nm 3 rad -1 kg.
- the gearbox support tilt stiffness to moment of inertia product may be in the range from 3.0x10 12 Nm 3 rad -1 kg to 9.0x10 15 Nm 3 rad -1 kg, and optionally in the range from 6.0x10 12 Nm 3 rad -1 kg to 4.0x10 15 Nm 3 rad -1 kg, and further optionally in the range from 2.5x10 13 Nm 3 rad -1 kg to 4.0x10 15 Nm 3 rad -1 kg.
- the gearbox support tilt stiffness to moment of inertia product may be in the range from 2.0x10 13 Nm 3 rad -1 kg to 4.0x10 15 Nm 3 rad -1 kg.
- the gearbox support tilt stiffness to moment of inertia product may be in the range from 2.5x10 13 Nm 3 rad -1 kg to 9.0x10 15 Nm 3 rad -1 kg.
- a product of the fan-gearbox axial distance multiplied by the moment of inertia of the fan may be defined.
- the product of the fan-gearbox axial distance and the moment of inertia of the fan may be greater than or equal to 1.9x10 7 kgm 3 , and optionally greater than or equal to 2.9x10 7 kgm 3 .
- the product of the fan-gearbox axial distance and the moment of inertia of the fan may be greater than or equal to 2.0x10 7 kgm 3 .
- the product of the fan-gearbox axial distance and the moment of inertia of the fan may be greater than or equal to 1.2x10 8 kgm 3 .
- the product of the fan-gearbox axial distance and the moment of inertia of the fan may be in the range from 1.9x10 7 kgm 3 to 6.2x10 8 kgm 3 , and optionally in the range from 2.9x10 7 kgm 3 to 3.9x10 8 kgm 3 .
- the product of the fan-gearbox axial distance and the moment of inertia of the fan may be in the range from 2.0x10 7 kgm 3 to 8.0x10 7 kgm 3 .
- the product of the fan-gearbox axial distance and the moment of inertia of the fan may be in the range from 1.2x10 8 kgm 3 to 6.2x10 8 kgm 3 .
- a ratio given by the fan-gearbox axial distance divided by the moment of inertia of the fan may also be defined.
- the fan-gearbox axial distance divided by the moment of inertia of the fan may be less than or equal to 8.8x10 -9 m/kgm 2 (i.e. kg -1 m -1 ), and optionally less than or equal to 6.2x10 -9 m/kgm 2 .
- the fan-gearbox axial distance divided by the moment of inertia of the fan may be less than or equal to 6.5x10 -9 m/kgm 2 or less than or equal to 5.2x10 -9 m/kgm 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan-gearbox axial distance divided by the moment of inertia of the fan may be less than or equal to 2.8x10 -9 m/kgm 2 or less than or equal to 1.8x10 -9 m/kgm 2 .
- the fan-gearbox axial distance divided by the moment of inertia of the fan may be in the range from 5.3x10 -10 m/kgm 2 to 8.8x10 -9 m/kgm 2 , and optionally in the range from 8.8x10 -10 m/kgm 2 to 6.2x10 -9 m/kgm 2 .
- the fan-gearbox axial distance divided by the moment of inertia of the fan may be in the range from 2.1x10 -9 m/kgm 2 to 6.5x10 -9 m/kgm 2 and optionally in the range from 4.0x10 -9 m/kgm 2 to 5.2x10 -9 m/kgm 2 (and may be equal to 4.6x10 -9 m/kgm 2 ).
- the fan-gearbox axial distance divided by the moment of inertia of the fan may be in the range from 5.3x10 -10 m/kgm 2 to 2.8x10 -9 m/kgm 2 , and optionally in the range from 8.0x10 -10 m/kgm 2 to 1.8x10 -9 m/kgm 2 (and may be equal to 1.3x10 -9 m/kgm 2 ).
- a first gearbox support strength ratio may be defined as the torsional strength of the gearbox support 40 radial bending stiffness of the gearbox support 40 ⁇ the cross sectional area of the gearbox
- the first gearbox support strength ratio may be greater than or equal to 7.0x10 -3 , and optionally greater than or equal to 1.0x10 -2 or greater than or equal to 2.0x10 -2 .
- the first gearbox support strength ratio may be in the range from 7.0x10 -3 to 2.5x10 -1 , and optionally in the range from 1.0x10 -2 to 1.0x10 -1 , in the range from 7.0x10 -3 to 2.0x10 -2 or in the range from 2.0x10 -2 to 2.5x10 -1 .
- the first gearbox support strength ratio may be in the range from 1.9x10 -2 to 2.9x10 -2 (and may be equal to 2.4x10 -2 ).
- a second gearbox support strength may be defined as: the torsional strength of the gearbox support 40 tilt stiffness of the gearbox support 40 ⁇ the planet gear spacing angle ⁇
- the second gearbox support strength ratio may be greater than or equal to 1.0x10 -1 , and optionally greater than or equal to 1.5x10 -1 , greater than or equal to 1.0x10 -1 or greater than or equal to 2.5x10 -1 .
- the second gearbox support strength ratio may be in the range from 1.0x10 -1 to 3.5, and optionally in the range from 1.5x10 -1 to 1.7, in the range from 1.0x10 -1 to 2.5x10 -1 or in the range from 2.5x10 -1 to 3.5.
- the second gearbox support strength ratio may be in the range from 2.8x10 -1 to 3.8x10 -1 (and may be equal to 3.3x10 -1 ).
- a first gearbox support shear stress ratio may be defined as: the torsional shear stress of the gearbox support 40 at MTO radial bending stiffness of the gearbox support 40
- the first gearbox support shear stress ratio may be less than or equal to 4.9x10 1 m -1 , and optionally less than or equal to 20 m -1 .
- the first gearbox support shear stress ratio may be less than or equal to 35 m -1 or less than or equal to 10.0 m -1 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first gearbox support shear stress ratio may be less than or equal to 12 m -1 or less than or equal to 6.0 m -1 .
- the first gearbox support shear stress ratio may be in the range from 0.35 m -1 to 4.9x10 1 m -1 , and optionally in the range from 0.70 m -1 to 20 m -1 .
- the first gearbox support shear stress ratio may be in the range from 0.70 m -1 to 35 m -1 and optionally in the range from 6.0 m -1 to 10.0 m -1 (and may be equal to 7.8 m -1 ).
- the first gearbox support shear stress ratio may be in the range from 0.50 m -1 to 12 m -1 , and optionally in the range from 3.0 m -1 to 6.0 m -1 (and may be equal to 5.4 m -1 ).
- a second gearbox support shear stress ratio may be defined as: the torsional shear stress of the gearbox support 40 at maximum take off conditions tilt stiffness of the gearbox support 40
- the second gearbox support shear stress ratio may be less than or equal to 4.1x10 3 rad/m 3 , and optionally less than or equal to 1.4x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be less than or equal to 2.9x10 3 rad/m 3 or less than or equal to 6.5x10 2 rad/m 3 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second gearbox support shear stress ratio may be less than or equal to 7.0x10 2 rad/m 3 or less than or equal to 2.5x10 2 rad/m 3 .
- the second gearbox support shear stress ratio may be in the range from 6.6 rad/m 3 to 4.1x10 3 rad/m 3 , and optionally in the range from 125x10 1 rad/m 3 to 1.4x10 3 rad/m 3 .
- the second gearbox support shear stress ratio may be in the range from 2.9x10 1 rad/m 3 to 2.9x10 3 rad/m 3 and optionally in the range from 2.5x10 2 rad/m 3 to 6.5x10 2 rad/m 3 (and may be equal to 4.5x10 2 rad/m 3 ).
- the second gearbox support shear stress ratio may be in the range from 1.0x10 1 rad/m 3 to 7.0x10 2 rad/m 3 , and optionally in the range from 50 rad/m 3 to 2.5x10 2 rad/m 3 (and may be equal to 1.5x10 2 rad/m 3 ).
- a flight cycle ratio may be defined as: the torsional shear stress of the gearbox support 40 at maximum take off conditions the torsional shear stress of the gearbox support 40 at cruise conditions
- the flight cycle ratio may be less than or equal to 3.20, and optionally less than or equal to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85, or optionally less than or equal to 2.75.
- the flight cycle ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the flight cycle ratio may be less than or equal to 2.8 or less than or equal to 2.7.
- the flight cycle ratio may be in the range from 2.10 to 3.20, and optionally in the range from 2.3 to 2.9.
- the flight cycle ratio may be in the range from 2.3 to 3.2 and optionally in the range from 2.8 to 3.0 (and may be equal to 2.9).
- the flight cycle ratio may be in the range from 2.1 to 2.8, and optionally in the range from 2.3 to 2.7 (and may be equal to 2.6).
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the flight cycle product) may be defined as: the torsional shear stress of the gearbox support 40 at maximum take off conditions ⁇ the torsional shear stress of the gearbox support 40 at cruise conditions
- the flight cycle product may be greater than or equal to 1.00x10 16 (N/m 2 ) 2 , and optionally greater than or equal to 2.05x10 16 (N/m 2 ) 2 .
- the flight cycle product may be greater than or equal to 2.0x10 16 (N/m 2 ) 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the flight cycle product may be greater than or equal to 2.2x10 16 (N/m 2 ) 2 .
- the flight cycle product may be in the range from 1.00x10 16 (N/m 2 ) 2 to 7.50x10 16 (N/m 2 ) 2 , and optionally in the range from 2.05x10 16 (N/m 2 ) 2 to 4.9x10 16 (N/m 2 ) 2 .
- the flight cycle product may be in the range from 2.0x10 16 (N/m 2 ) 2 to 4.9x10 16 (N/m 2 ) 2 .
- the flight cycle product may be in the range from 2.2x10 16 (N/m 2 ) 2 to 6.2x10 16 (N/m 2 ) 2 .
- a first torque transmission ratio may be defined as: the torque transmitted through the gearbox 30 at maximum takeoff conditions the torque transmitted through the gearbox 30 at cruise conditions
- the first torque transmission ratio may be less than or equal to 3.2, and optionally less than or equal to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85 or optionally less than or equal to 2.75.
- the first torque transmission ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first torque transmission ratio may be less than or equal to 2.8 or less than or equal to 2.7.
- the first torque transmission ratio may be in the range from 2.1 to 3.2, and optionally in the range from 2.3 to 2.9.
- the first torque transmission ratio may be in the range from 2.3 to 3.2 and optionally in the range from 2.8 to 3.0 (and may be equal to 2.9).
- the first torque transmission ratio may be in the range from 2.1 to 2.8, and optionally in the range from 2.5 to 2.7 (and may be equal to 2.6).
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the first torque transmission product) may be defined as: the torque transmitted through the gearbox 30 at maximum take off conditions ⁇ the torque transmitted through the gearbox 30 at cruise conditions
- the first torque transmission product may be greater than or equal to 2.1x10 9 (Nm) 2 , and optionally greater than or equal to 3.5x10 9 (Nm) 2 .
- the first torque transmission product may be greater than or equal to 4.0x10 9 (Nm) 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first torque transmission product may be greater than or equal to 9.0x10 9 (Nm) 2 .
- the first torque transmission product may be in the range from 2.1x10 9 (Nm) 2 to 9.0x10 10 (Nm) 2 , and optionally in the range from 3.5x10 9 (Nm) 2 to 5.2x10 10 (Nm) 2 .
- the first torque transmission product may be in the range from 4.0x10 9 (Nm) 2 to 9.0x10 9 (Nm) 2 .
- the first torque transmission product may be in the range from 9.0x10 9 (Nm) 2 to 9.0x10 10 (Nm) 2 .
- a second torque transmission ratio may be defined as: the torque transmitted through the gearbox support 40 at maximum take off conditions the torque transmitted through the gearbox support 40 at cruise conditions
- the second torque transmission ratio may be less than or equal to 3.2, and optionally less than or equal to 2.95, or optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85, or optionally less than or equal to 2.75.
- the second torque transmission ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second torque transmission ratio may be greater than or equal to 2.8 or less than or equal to 2.7.
- the second torque transmission ratio may be in the range from 2.1 to 3.2, and optionally in the range from 2.3 to 2.9.
- the second torque transmission ratio may be in the range from 2.3 to 3.2 and optionally in the range from 2.8 to 3.0 (and may be equal to 2.9).
- the second torque transmission ratio may be in the range from 2.1 to 2.8, and optionally in the range from 2.3 to 2.7 (and may be equal to 2.6).
- a product of the parameters making up that ratio may be defined.
- This product (referred to as the second torque transmission product) may be defined as: the torque transmitted through the gearbox support 40 at maximum take off conditions ⁇ the torque transmitted through the gerbox support 40 at cruise conditions
- the second torque transmission product may be greater than or equal to 4.1x10 9 (Nm) 2 , and optionally greater than or equal to 6.1x10 9 (Nm) 2 .
- the second torque transmission product may be greater than or equal to 4.1x10 9 (Nm) 2 . In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second torque transmission product may be greater than or equal to 1.1x10 10 (Nm) 2 .
- the second torque transmission product may be in the range from 4.1x10 9 (Nm) 2 to 9.0x10 10 (Nm) 2 , and optionally in the range from 6.1x10 9 (Nm) 2 to 8.0x10 10 (Nm) 2 .
- the second torque transmission product may be in the range from 4.1x10 9 (Nm) 2 to 3.9x10 10 .
- the second torque transmission product may be in the range from 1.1x10 10 (Nm) 2 to 9.0x10 10 (Nm) 2 .
- Figure 19 illustrates an example aircraft 1000 having a gas turbine engine 10 attached to each wing 1002a, 1002b thereof. Each gas turbine engine 10 is attached via a respective pylon 1004a, 1004b. The gas turbines 10 may be that of any embodiment described herein.
- the aircraft shown in Figure 19 is to be understood as the aircraft for which the gas turbine engine 10 of any embodiment or aspect disclosed herein has been designed to be attached.
- the aircraft 1000 has a cruise condition corresponding to the cruise conditions defined elsewhere herein and a max take-off condition corresponding to the MTO conditions defined elsewhere herein.
- the present disclosure also relates to a method 2000 of operating a gas turbine engine on an aircraft (e.g. the aircraft of Figure 19 ).
- the method 2000 is illustrated in Figure 20 .
- the method 2000 comprises operating 2010 the gas turbine engine 10 described elsewhere herein to provide propulsion for the aircraft to which it is mounted under maximum take-off conditions.
- the method further comprises operating 2020 the gas turbine engine to provide propulsion during cruise conditions.
- the gas turbine engine is operated such that any of the parameters or ratios defined herein are within the specified ranges. Cruise conditions and max-take off conditions are as defined elsewhere herein.
- the torque on the core shaft 26 may be referred to as the input torque, as this is the torque which is input to the gearbox 30.
- the torque supplied by the turbine 19 to the core shaft (i.e. the torque on the core shaft) at cruise conditions may be greater than or equal to 10,000 Nm, and optionally greater than or equal to 11,000 Nm.
- the torque on the core shaft 26 at cruise conditions may be greater than or equal to 10,000 or 11,000 Nm (and optionally may be equal to 12,760 Nm).
- the torque on the core shaft 26 at cruise conditions may be greater than or equal to 25,000 Nm, and optionally greater than or equal to 30,000 Nm (and optionally may be equal to 34,000 Nm).
- the torque on the core shaft at cruise conditions may be in the range from 10,000 to 50,000 Nm, and optionally from 11,000 to 45,000 Nm.
- the torque on the core shaft 26 at cruise conditions may be in the range from 10,000 to 15,000 Nm, and optionally from 11,000 to 14,000 Nm (and optionally may be equal to 12,760 Nm).
- the torque on the core shaft 26 at cruise conditions may be in the range from 25,000 Nm to 50,000 Nm, and optionally from 30,000 to 40,000 Nm (and optionally may be equal to 34,000 Nm).
- the torque on the core shaft 26 may be greater than or equal to 28,000 Nm, and optionally greater than or equal to 30,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque on the core shaft 26 under MTO conditions may be greater than or equal to 28,000, and optionally greater than or equal to 35,000 Nm (and optionally may be equal to 36,300 Nm).
- the torque on the core shaft 26 under MTO conditions may greater than or equal to 70,000 Nm, and optionally greater than or equal to 80,000 or 82,000 Nm (and optionally may be equal to 87,000 Nm).
- the torque on the core shaft 26 may be in the range from 28,000 Nm to 135,000 Nm, and optionally in the range from 30,000 to 110,000 Nm.
- the torque on the core shaft 26 under MTO conditions may be in the range from 28,000 to 50,000 Nm, and optionally from 35,000 to 38,000 Nm (and optionally may be equal to 36,300 Nm).
- the torque on the core shaft 26 under MTO conditions may be in the range from 70,000 Nm to 135,000 Nm, and optionally from 80,000 to 90,000 Nm or 82,000 to 92,000 Nm (and optionally may be equal to 87,000 Nm).
- Torque has units of [force]x[distance] and may be expressed in units of Newton metres (N.m), and is defined in the usual way as would be understood by the skilled person.
- Figure 21 illustrates how the stiffnesses defined herein may be measured.
- Figure 21 shows a plot of the displacement ⁇ resulting from the application of a load L (e.g. a force, moment or torque) applied to a component for which the stiffness is being measured.
- L e.g. a force, moment or torque
- L S a load, moment or torque
- the stiffnesses defined herein may be determined by measuring the gradient of the linear region between points R and S (with the stiffness being the inverse of that gradient).
- the gradient may be found for as large a region of the linear region as possible to increase the accuracy of the measurement by providing a larger displacement to measure.
- the gradient may be found by applying a load equal to or just greater than L R and equal to or just less than L S .
- the displacement is referred to as ⁇ in this description, the skilled person would appreciate that equivalent principles would apply to a linear or angular displacement.
- stiffnesses defined herein are for the corresponding component(s) when the engine is off (i.e. at zero speed/on the bench).
- the stiffness generally does not vary over the operating range of the engine; the stiffness at cruise conditions of the aircraft to which the engine is used (those cruise conditions being as defined elsewhere herein) may therefore be the same as for when the engine is not in use.
- the stiffnesses defined herein are to be understood as being values for when the engine is at room temperature and unmoving. Values for component strength (e.g. the torsional strength of the gearbox support) as given herein are also at room temperature unless otherwise stated.
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Abstract
Description
- The present disclosure relates to gas turbine engines, specifically gas turbine engines for aircraft. Aspects of the present disclosure also relate to an aircraft comprising the gas turbine engine, and a method of operating the gas turbine engine.
- Gas turbine engines for aircraft propulsion have many design factors that affect the overall efficiency and power output or thrust. A general aim for a gas turbine engine is to provide low specific fuel consumption (SFC). To enable a higher thrust at a high efficiency, a larger diameter fan may be used. In order to facilitate use of a larger fan size, a gearbox is provided having an output to a fan shaft via which the fan is driven. The gearbox receives drive from a core shaft connected to a turbine system of the engine core. The gearbox allows the fan to operate at a reduced rotational speed compared to if a direct drive were used.
- When making an engine having a larger fan diameter however, the inventor has discovered that simply scaling up components of a known engine type may not lead to an efficient design. For example, there may be problems associated with mounting the gearbox and the fan shaft within the engine. Consideration of the properties of components used to mount the gearbox and the properties of the fan shaft are therefore required.
- According to a first aspect there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a radial bending stiffness to moment of inertia ratio of:
- The radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg-1m-3. The radial bending stiffness to moment of inertia ratio may be in the range from 2.5x10-2 Nkg-1m-3 to 6.0 Nkg-1m-3. The radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg-1m-3 to 3.0 Nkg-1m-3. The radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg-1m-3 to 0.6 Nkg-1m-3
- A ratio of the radial bending stiffness of the fan shaft at the output of the gearbox to the moment of inertia of the fan may be greater than or equal to 2.5x10-2 Nkg-1m-3, greater than or equal to 0.05 Nkg-1m-3, in the range from 2.5x10-2 Nkg-1m-3 to 6.0 Nkg-1m-3, in the range from 0.05 Nkg-1m-3 to 3.0 Nkg-1m-3, or in the range from 0.05 Nkg-1m-3 to 0.6 Nkg-1m-3.
- A ratio of the radial bending stiffness of the gearbox support to the moment of inertia of the fan may be greater than or equal to 3.0x10-2 Nkg-1m-3, greater than or equal to 0.06 Nkg-1m-3, in the range from 3.0x10-2 Nkg-1m-3 to 4.0 Nkg-1m-3, in the range from 0.06 Nkg-1m-3 to 2.0 Nkg-1m-3, or in the range from 0.06 Nkg-1m-3 to 0.48 Nkg-1m-3.
- The radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 4.00x106 N/m. The radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 3.7x107 N/m. The radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 4.00x106 N/m to 1.50x109 N/m. The radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 3.7x107 N/m to 1.0x109 N/m.
- The radial bending stiffness of the gearbox support may be greater than or equal to 1.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 2.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 3.0x107 N/m. The radial bending stiffness of the gearbox support may be in the range from 1.0x107 N/m to 4.0x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 2.0x107 N/m to 3.0x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 3.0x107 N/m to 2.0x108 N/m.
- The diameter of the fan may be in the range from 240 cm to 280 cm. In such an embodiment, the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg-1m-3 or in the range from 0.05 Nkg-1m-3 to 4.0 Nkg-1m-3.
- Alternatively, the diameter of the fan may be in the range from 330 cm to 380 cm. In such an embodiment, the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.025 Nkg-1m-3 or in the range from 0.025 Nkg-1m-3 to 2.0 Nkg-1m-3.
- A product (e.g. a radial bending stiffness moment of inertia product) of:
- A tilt stiffness to moment of inertia ratio of:
- According to a second aspect there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a tilt stiffness to moment of inertia ratio of:
- The tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10-3 Nrad-1kg-1m-1. The tilt stiffness to moment of inertia ratio may be in the range from 4.0x10-4 Nrad-1kg-1m-1 to 2.7x10-1 Nrad-1kg-1m-1. The tilt stiffness to moment of inertia ratio may be in the range from 1.0x10-3 Nrad-1kg-1m-1 to 0.1 Nrad-1kg-1m-1. The tilt stiffness to moment of inertia ratio may be in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.5x10-2 Nrad-1kg-1m-1.
- A ratio of the tilt stiffness of the fan shaft at the output of the gearbox to the moment of inertia of the fan may be greater than or equal to 4.0x10-4 Nrad-1kg-1m-1, greater than or equal to 1.0x10-3 Nrad-1kg-1m-1, in the range from 4.0x10-4 Nrad-1kg-1m-1 to 2.7x10-1 Nrad-1kg-1m-1, in the range from 1.0x10-3 Nrad-1kg-1m-1 to 0.1 Nrad-1kg-1m-1, or in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.5x10-2 Nrad-1kg-1m-1.
- A ratio of the tilt stiffness of the gearbox support to the moment of inertia of the fan may be greater than or equal to 1.0x10-3 Nrad-1kg-1m-1, greater than or equal to 2.0x10-3 Nrad-1kg-1m-1, in the range from 1.0x10-3 Nrad- 1kg-1m-1 to 7.0x10-2 Nrad-1kg-1m-1, in the range from 2.0x10-3 Nrad-1kg-1m-1 to 3.0x10-2 Nrad-1kg-1m-1, or in the range from 2.0x10-3 Nrad-1kg-1m-1 to 7.0x10-3 Nrad-1kg-1m-1.
- The tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 7.00x104 Nm/rad. The tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 9.5x105 Nm/rad. The tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 7.00x104 Nm/rad to 7.00x107 Nm/rad. The tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 9.5x105 Nm/rad to 3.5x107 Nm/rad.
- The tilt stiffness of the gearbox support may be greater than or equal to 1.2x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 2.4x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 3.9x105 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 1.2x105 Nm/rad to 2.1x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 2.4x105 Nm/rad to 1.6x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 3.9x105 Nm/rad to 9.0x106 Nm/rad.
- A radial bending stiffness to moment of inertia ratio of:
- The diameter of the fan may be in the range from 240 cm to 28 cm. In such an example, the tilt stiffness to moment of inertia ratio greater may be than or equal to 1.0x10-3 Nrad-1kg-1m-1 or in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.45x10-2 Nrad-1kg-1m-1.
- Alternatively, the diameter of the fan may be in the range from 330 cm to 380 cm. In such an embodiment the tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10-4 Nrad-1kg-1m-1 or in the range from 4.0x10-4 Nrad-1kg-1m-1 to 3.0x10-2 Nrad-1kg-1m-1.
- A product (e.g. the tilt stiffness moment of inertia product) of:
- One or more of the following features may apply to either or both of the first and second aspects above:
The moment of inertia of the fan may be greater than or equal to 8.3x107 kgm2. The moment of inertia of the fan may be in the range from 7.40x107 kgm2 to 9.00x108 kgm2. The moment of inertia of the fan may be in the range from 8.3x107 kgm2 to 6.5x108 kgm2. - The fan blades may be formed from a metallic material. The fan blades may be formed from titanium or an aluminium-lithium alloy with a titanium leading edge.
- The fan blades may be formed at least partly from an organic matrix composite. The fan blades may have a carbon composite body with a metallic leading edge.
- The fan shaft may connect the output of the gearbox to the fan. A gearbox output position may be defined as the point of connection between the fan shaft and the gearbox. The fan may have an axial centreline. A fan-gearbox axial distance may be defined as the axial distance between the axial position of the gearbox output position and the fan axial centreline. The fan-gearbox axial distance multiplied by the moment of inertia of the fan may be greater than or equal to 1.9x107 kgm3, greater than or equal to 2.9x107 kgm3, in the range from 1.9x107 kgm3 to 6.2x108 kgm3, or in the range from 2.9x107 kgm3 to 3.9x108 kgm3.
- The fan shaft may be defined as the torque transfer component extending from the output of the gearbox to the input to the fan. The fan shaft may comprise at least part of a gearbox output shaft and at least part of a fan input shaft.
- The gearbox may be in a star configuration and the output of the gearbox may be a gearbox output position defined as the point of connection between the ring gear and the fan shaft. Alternatively, the gearbox may be in a planetary configuration and the output of the gearbox may be a gearbox output position at the interface between the fan shaft and the planet carrier.
- According to a third aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; a gearbox support arranged to at least partially support the gearbox within the propulsor, and wherein: the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a radial bending stiffness to moment of inertia ratio of:
- The propulsor of the third aspect may have some or all of the features described above with respect to the gas turbine engine of the first aspect, and may be a gas turbine engine in some embodiments.
- According to a fourth aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a gearbox support arranged to at least partially support the gearbox within the propulsor, and wherein: the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a tilt stiffness to moment of inertia ratio of:
- The propulsor of the fourth aspect may have some or all of the features described above with respect to the gas turbine engine of the second aspect, and may be a gas turbine engine in some embodiments.
- The third and fourth aspects may be combined. According to a fifth aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a gearbox support arranged to at least partially support the gearbox within the propulsor,
and wherein the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and wherein: - a) a radial bending stiffness to moment of inertia ratio of:
- b) a tilt stiffness to moment of inertia ratio of:
- The propulsor of the fifth aspect may have some or all of the features described above with respect to the gas turbine engine of the first or second aspect, and may be a gas turbine engine in some embodiments.
- According to a sixth aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a radial bending stiffness to moment of inertia ratio of:
- The method of the sixth aspect may be a method of operating the gas turbine engine or the propulsor of the first aspect or third aspect respectively. Any of the features, ratios and parameters introduced above in connection with the first or third aspect may also therefore apply to the sixth aspect
- According to a seventh aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine,
wherein the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a tilt stiffness to moment of inertia ratio of: - The method of the seventh aspect may be a method of operating the gas turbine engine or the propulsor of the second aspect or fourth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the second aspect or fourth aspect may also therefore apply to the seventh aspect.
- The sixth and seventh aspects may be combined. According to an eighth aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2, and wherein: a radial bending stiffness to moment of inertia ratio of:
b) a tilt stiffness to moment of inertia ratio of: - The method of the eighth aspect may be a method of operating the gas turbine engine or the propulsor of the first, second or fifth aspect. Any of the features, ratios and parameters introduced above in connection with the first, second or fifth aspect may also therefore apply to the eighth aspect.
- The inventor has discovered that either the radial bending/tilt stiffness of one or both of the fan shaft at the gearbox output and the gearbox support need to be low enough for a given moment of inertia of the fan to isolate the gearbox from the transmission of loads from the fan. Rotation of the fan will cause a gyroscopic effect meaning that the fan shaft will tend to maintain a steady direction of its axis of rotation. During manoeuvring of the aircraft to which the gas turbine engine is mounted the orientation of the axis of rotation of the fan shaft will however change. The inventor has found that gyroscopic effects will result in loads being transmitted to the gearbox. The inventor has discovered that by using a fan shaft and/or gearbox support having a radial bending/tilt stiffness so that the radial bending/stilt stiffness to moment of inertia ratio defined above is within the specified range the problem of load from the fan being transmitted to the gearbox can be addressed. The inventor has found that if the radial bending/tilt stiffness of the fan shaft or the gearbox support shaft is increased so that the ratio is outside of the defined range there is inadequate isolation of the gearbox. The inventor has also found that if the fan shaft/gearbox support stiffness were to be decreased so that the ratio is outside of the range above excessive vibration at low modal frequencies would result.
- In other aspects, value ranges for the product of the components of the radial bending stiffness to moment of inertia ratio and the tilt stiffness to moment of inertia ratio may be specified instead of, or as well as, value ranges for the ratios.
- According to one such aspect, the first aspect introduced above may be formulated as an aspect providing a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a product of:
- According to another such aspect, the second aspect introduced above may be formulated as an aspect providing a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the moment of inertia of the fan is greater than or equal to 7.40x107 kgm2; and a product of:
- The skilled person will appreciate that method and propulsor aspects may be formulated accordingly.
- According to a ninth aspect there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2; and a first gearbox support strength ratio of:
- The first gearbox support strength ratio may be greater than or equal to 1.0x10-2. The first gearbox support strength ratio may be greater than or equal to 7.0x10-3. The first gearbox support strength ratio may be greater than or equal to 2.0x10-2. The first gearbox support strength ratio may be in the range from 7.0x10-3 to 2.5x10-1. The first gearbox support strength ratio may be in the range from 1.0x10-2 to 1.0x10-1. The first gearbox support strength ratio may be in the range from 7.0x10-3 to 2.0x10-2. The first gearbox support strength ratio may be in the range from 2.0x10-2 to 2.5x10-1.
- The radial bending stiffness of the gearbox support may be greater than or equal to 1.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 2.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 3.0x107 N/m. The radial bending stiffness of the gearbox support may be in the range from 1.0x107 N/m to 4.0x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 2.0x107 N/m to 3.0x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 3.0x107 N/m to 2.0x108 N/m.
- The torsional strength of the gearbox support may be greater than or equal to 1.60x105 Nm. The torsional strength of the gearbox support may be greater than or equal to 1.8x105 Nm. The torsional strength of the gearbox support may be in the range from 1.60x105 Nm to 2.00x107 Nm. The torsional strength of the gearbox support may be in the range from 1.8x105 Nm to 1.5x106 Nm.
- A planet gear spacing angle in radians may be defined as 2π/N, where N is the number of planet gears.
- A second gearbox support strength ratio of:
- According to a tenth aspect there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: a planet gear spacing angle in radians is defined as 2π/N, where N is the number of planet gears; the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2; and a second gearbox support strength ratio of:
- The second gearbox support strength ratio may be greater than or equal to 1.5x10-1. The second gearbox support strength ratio may be greater than or equal to 2.5x10-1. The second gearbox support strength ratio may be in the range from 1.0x10-1 to 3.5. The second gearbox support strength ratio may be in the range from 1.5x10-1 to 1.7. The second gearbox support strength ratio may be in the range from 1.0x10-1 to 2.5x10-1. The second gearbox support strength ratio may be in the range from 2.5x10-1 to 3.5.
- The tilt stiffness of the gearbox support may be greater than or equal to 1.2x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 2.4x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 3.9x105 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 1.2x105 Nm/rad to 2.1x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 2.4x105 Nm/rad to 1.6x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 3.9x105 Nm/rad to 9.0x106 Nm/rad.
- The torsional strength of the gearbox support may be greater than or equal to 1.60x105 Nm. The torsional strength of the gearbox support may be greater than or equal to 1.8x105 Nm. The torsional strength of the gearbox support may be in the range from 1.60x105 Nm to 2.00x107 Nm. The torsional strength of the gearbox support may be in the range from 1.8x105 Nm to 1.5x106 Nm.
- A first gearbox support strength ratio of:
- One or more of the following features may apply to either or both of the ninth and tenth aspects above:
The gearbox may be in a star configuration. - The planet gear spacing angle may be greater than or equal to 9.0x10-1 rad. The planet gear spacing angle may be in the range between 9.0x10-1 rad to 2.1 rad.
- The cross sectional area of the gearbox may be greater than or equal to 2.6x10-1 m2. The cross sectional area of the gearbox may be in the range from 2.4x10-1 m2 to 1.10 m2. The cross sectional area of the gearbox may be in the range from 2.6x10-1 m2 to 9.0x10-1 m2. The cross sectional area (CSA) of the gearbox may be defined as the area of the pitch circle of the ring gear.
- A torsional shear stress of the gearbox support, at maximum take-off conditions, may be less than or equal to 4.90x108 N/m2, less than or equal to 3.5x108 N/m2, in the range from 1.40x108 N/m2 to 4.90x108 N/m2, or in the range from 2.0x108 N/m2 to 3.5x108 N/m2.
-
-
- According to an eleventh aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2; a first gearbox support strength ratio of:
- The propulsor of the eleventh aspect may have some or all of the features described above with respect to the gas turbine engine of the ninth aspect, and may be a gas turbine engine in some embodiments.
- According to an twelfth aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2; and a planet gear spacing angle in radians is defined as 2π/N, where N is the number of planet gears; and a second gearbox support strength ratio of:
- The propulsor of the twelfth aspect may have some or all of the features described above with respect to the gas turbine engine of the tenth aspect, and may be a gas turbine engine in some embodiments.
- The eleventh and twelfth aspects may be combined. According to a thirteenth aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2, and wherein: a first gearbox support strength ratio of:
b) a planet gear spacing angle in radians is defined as 2π/N, where N is the number of planet gears; and a second gearbox support strength ratio of: - The propulsor of the thirteenth aspect may have some or all of the features described above with respect to the gas turbine engine of the ninth aspect and tenth aspects, and may be a gas turbine engine in some embodiments.
- According to a fourteenth aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2; and a first gearbox support strength ratio of:
- The method of the fourteenth aspect may be a method of operating the gas turbine engine or the propulsor of the ninth aspect or eleventh aspect respectively. Any of the features, ratios and parameters introduced above in connection with the ninth aspect or eleventh aspect may also therefore apply to the fourteenth aspect.
- According to a fifteenth aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2; and a planet gear spacing angle in radians is defined as 2π/N, where N is the number of planet gears; and a second gearbox support strength ratio of:
- The method of the fifteenth aspect may be a method of operating the gas turbine engine or the propulsor of the tenth aspect or twelfth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the tenth aspect or twelfth aspect may also therefore apply to the fifteenth aspect.
- The fourteenth and fifteenth aspects may be combined. According to a sixteenth aspect there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein the gearbox has a cross sectional area, the cross sectional area being greater than or equal to 2.4x10-1 m2, and
- a) a first gearbox support strength ratio of:
- b) a planet gear spacing angle in radians is defined as 2π/N, where N is the number of planet gears; and a second gearbox support strength ratio of:
- The method of the sixteenth aspect may be a method of operating the gas turbine engine or the propulsor of the ninth, tenth or thirteenth aspect. Any of the features, ratios and parameters introduced above in connection with the ninth, tenth or thirteenth aspect may also therefore apply to the sixteenth aspect.
- The inventor has discovered that by designing the gearbox support so that the ratio of its torsional strength to its radial bending stiffness and cross sectional area (i.e. the first gearbox support strength ratio) is within the specified range sufficient strength of the support is provided so that the engine is reliable with sufficient stiffness provided to minimise misalignment of the gears in the gearbox and avoid vibration. The inventor has found that a similar consideration applies to the ratio of the torsional strength to the tilt stiffness and planet gear spacing angle (i.e. the second gearbox support strength ratio).
- According to a seventeenth aspect there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, wherein: a first gearbox support shear stress ratio of:
- The first gearbox support shear stress ratio may be less than or equal to 20 m-1. The first gearbox support shear stress ratio may be in the range from 3.5x10-1 m-1 to 4.9x101 m-1. The first gearbox support shear stress ratio may be in the range from 0.70 m-1 to 20 m-1.
- The radial bending stiffness of the gearbox support may be greater than or equal to 1.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 2.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 3.0x107 N/m. The radial bending stiffness of the gearbox support may be in the range from 1.0x107 N/m to 4.0x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 2.0x107 N/m to 3x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 3.0x107 N/m to 2.0x108 N/m.
- The torsional shear stress of the gearbox support, at maximum take-off conditions, may be less than or equal to 4.90x108 N/m2. The torsional shear stress of the gearbox support, at maximum take-off conditions, may be less than or equal to 3.5x108 N/m2. The torsional shear stress of the gearbox support, at maximum take-off conditions, may be in the range from 1.40x108 N/m2 to 4.90x108 N/m2. The torsional shear stress of the gearbox support, at maximum take-off conditions, may be in the range from 2.0x108 N/m2 to 3.5x108 N/m2.
- The diameter of the fan may be in the range from 240 cm to 280 cm. In such an embodiment, the first gearbox support shear stress ratio may be less than or equal to 35 m-1, or in the range from 0.70 m-1 to 35 m-1. Alternatively, the diameter of the fan may be in the range from 330 cm to 380 cm. In such an embodiment the first gearbox support shear stress ratio may be less than or equal to 12 m-1, or in the range from 0.50 m-1 to 12 m-1.
- A second gearbox support shear stress ratio of:
- According to an eighteenth aspect there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox arranged to at least partially support the gearbox within the engine, and: a second gearbox support shear stress ratio of:
- The second gearbox support shear stress ratio may be less than or equal to 1.4x103 rad/m3. The second gearbox support shear stress ratio may be in the range from 6.6 rad/m3 to 4.1x103 rad/m3. The second gearbox support shear stress ratio may be in the range from 1.25x101 rad/m3 to 1.4x103 rad/m3.
- The tilt stiffness of the gearbox support may be greater than or equal to 1.2x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 2.4x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 3.9x105 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 1.2x105 Nm/rad to 2.1x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 2.4x105 Nm/rad to 1.6x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 3.9x105 Nm/rad to 9.0x106 Nm/rad.
- The torsional shear stress of the gearbox support, at maximum take-off conditions, may be less than or equal to 4.90x108 N/m2. The torsional shear stress of the gearbox support, at maximum take-off conditions, may be less than or equal to 3.5x108 N/m2. The torsional shear stress of the gearbox support, at maximum take-off conditions, may be in the range from 1.40x108 N/m2 to 4.90x108 N/m2. The torsional shear stress of the gearbox support, at maximum take-off conditions, may be in the range from 2.0x108 N/m2 to 3.5x108 N/m2.
- The diameter of the fan may be in the range from 240 cm to 280 cm. In such an embodiment, the second gearbox support shear stress ratio may be less than or equal to 2.9x103 rad/m3, or in the range from 2.9x101 rad/m3 to 2.9x103 rad/m3.
- Alternatively, the diameter of the fan may be in the range from 330 cm to 380 cm. In such an embodiment, the second gearbox support shear stress ratio may be less than or equal to 7.0x102 rad/m3 or in the range from 1.0x101 rad/m3 to 7.0x102 rad/m3.
- A first gearbox support shear stress ratio of:
- One or more of the following features may apply to either or both of the seventeenth and eighteenth aspects above:
The torsional strength of the gearbox support may be greater than or equal to 1.60x105 Nm. The torsional strength of the gearbox support may be greater than or equal to 1.8x105 Nm. The torsional strength of the gearbox support may be in the range from 1.60x105 Nm to 2.00x107 Nm. The torsional strength of the gearbox support may be in the range from 1.8x105 Nm to 1.5x106 Nm. - The cross sectional area of the gearbox may be greater than or equal to 2.4x10-1 m2. The cross sectional area of the gearbox may be greater than or equal to 2.6x10-1 m2. The cross sectional area of the gearbox may be in the range from 2.4x10-1 m2 to 1.10 m2. The cross sectional area of the gearbox may be in the range from 2.6x10-1 m2 to 9.0x10-1 m2.
- A planet gear spacing angle in radians may be defined as 2π/N, where N is the number of planet gears. The planet gear spacing angle may be greater than or equal to 9.0x10-1 rad. The planet gear spacing angle may be in the range between 9.0x10-1 rad to 2.1 rad.
- A first gearbox support strength ratio of:
- A second gearbox support strength ratio of:
- The torque transmitted through the gearbox support, at maximum take-off conditions, may be greater than or equal to 6.00x104 Nm, greater than or equal to 7.2x104 Nm, in the range from 6.00x104 Nm to 5.00x105 Nm, or in the range from 7.2x104 Nm to 4.2x105 Nm.
- The gearbox may be in a planetary configuration. In such an embodiment, the torque transmitted through the gearbox support, at maximum take-off conditions, may be greater than or equal to 6.00x104 Nm, greater than or equal to 7.2x104 Nm, in the range from 6.00x104 Nm to 3.00x105 Nm, or in the range from 7.2x104 Nm to 2.6x105 Nm.
- Alternatively, the gearbox may be in a star configuration. In such an embodiment the torque transmitted through the gearbox support, at maximum take-off conditions, may be greater than or equal to 1.10x105 Nm, greater than or equal to 1.3x105 Nm, in the range from 1.10x105 Nm to 5.00x105 Nm, or in the range from 1.3x105 Nm to 4.2x105 Nm.
- The maximum take-off conditions may be as defined anywhere herein, for example may be defined as operating at maximum take-off thrust for the engine at ISA sea level pressure and temperature +15°C with a fan inlet velocity of between 0.25 and 0.27 Mn, and optionally at 0.25 Mn.
- According to a nineteenth aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: a first gearbox support shear stress ratio of:
- The propulsor of the nineteenth aspect may have some or all of the features described above with respect to the gas turbine engine of the seventeenth aspect, and may be a gas turbine engine in some embodiments.
- According to a twentieth aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein: a second gearbox support shear stress ratio of:
- The propulsor of the twentieth aspect may have some or all of the features described above with respect to the gas turbine engine of the eighteenth aspect, and may be a gas turbine engine in some embodiments.
- The nineteenth and twentieth aspects may be combined. According to a twenty-first aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein:
- a) a first gearbox support shear stress ratio of:
- b) a second gearbox support shear stress ratio of:
- The propulsor of the twenty-first aspect may have some or all of the features described above with respect to the gas turbine engine of the seventeenth aspect and eighteenth aspects, and may be a gas turbine engine in some embodiments.
- According to a twenty-second aspect there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that, at maximum take-off conditions, a first gearbox support shear stress ratio of:
- The method of the twenty-second aspect may be a method of operating the gas turbine engine or the propulsor of the seventeenth aspect or the nineteenth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the seventeenth aspect or nineteenth aspect may also therefore apply to the twenty-second aspect.
- According to a twenty-third aspect there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that, at maximum take-off conditions, a second gearbox support shear stress ratio of:
- The method of the twenty-third aspect may be a method of operating the gas turbine engine or the propulsor of the eighteenth aspect or the twentieth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the eighteenth aspect or twentieth aspect may also therefore apply to the twenty-second aspect.
- In another aspect, the twenty-second and twenty-third aspects may be combined. In such an aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises: operating the gas turbine engine such that, at maximum take-off conditions:
- a) a first gearbox support shear stress ratio of:
- b) a second gearbox support shear stress ratio of:
- The method of the previous aspect may be a method of operating the gas turbine engine or the propulsor of the seventeenth, eighteenth or twenty-first aspect. Any of the features, ratios and parameters introduced above in connection with the seventeenth, eighteenth or twenty-first aspect may also therefore apply to this aspect.
- The inventor has discovered that by designing the gearbox support so that the ratio of its torsional shear stress at max take-off to its stiffness (radial bending or tilt stiffness) is within the specified range sufficient margin of strength of the support is provided at the highest load point in the engine operating cycle while the stiffness is suitable to ensure a suitable load share factor.
- According to a twenty-fourth aspect there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein:
a flight cycle ratio of: - The flight cycle ratio may be less than or equal to 2.95. The flight cycle ratio may be less than or equal to 2.9. The flight cycle ratio may be less than or equal to 2.90. The flight cycle ratio may be less than or equal to 2.85. The flight cycle ratio may be less than or equal to 2.75. The flight cycle ratio may be in the range from 2.10 to 3.20. The flight cycle ratio may be in the range from 2.3 to 2.9.
- The torsional shear stress of the gearbox support at maximum take-off conditions may be less than or equal to 4.90x108 N/m2. The torsional shear stress of the gearbox support at maximum take-off conditions may be less than or equal to 2.0x108 N/m2. The torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 1.40x108 N/m2 to 4.9x108 N/m2. The torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 2.0x108 N/m2 to 3.5x108 N/m2.
- The torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 7.00x107 N/m2. The torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 8.2x107 N/m2. The torsional shear stress of the gearbox support at cruise conditions may be in the range from 7.00x107 N/m2 to 1.90x108 N/m2. The torsional shear stress of the gearbox support at cruise conditions may be in the range from 8.2x107 N/m2 to 1.5x108 N/m2.
- The diameter of the fan may be in the range from 240 cm to 280 cm. In such an embodiment, the flight cycle ratio may be less than or equal to 3.2 or in the range from 2.3 to 3.2.
- Alternatively, the diameter of the fan may be in the range from 330 cm to 380 cm. In such an embodiment the flight cycle ratio may be less than or equal to 2.8 or in the range from 2.1 to 2.8.
- A product of:
- A first torque transmission ratio of:
- A second torque transmission ratio of:
- The torsional strength of the gearbox support may be greater than or equal to 1.60x105 Nm. The torsional strength of the gearbox support may be greater than or equal to 1.8x105 Nm. The torsional strength of the gearbox support may be in the range from 1.60x105 Nm to 2.00x107 Nm. The torsional strength of the gearbox support may be in the range from 1.8x105 Nm to 1.5x106 Nm.
- The gearbox may have a cross sectional area, and a first gearbox support strength ratio of:
- A planet gear spacing angle in radians may be defined as 2π/N, where N is the number of planet gears. A second gearbox support strength ratio of:
- The cross sectional area of the gearbox may be greater than or equal to 2.4x10-1 m2. The cross sectional area of the gearbox may be greater than or equal to 2.6x10-1 m2. The cross sectional area of the gearbox may be in the range from 2.4x10-1 m2 to 1.10 m2. The cross sectional area of the gearbox may be in the range from 2.6x10-1 m2 to 9.0x10-1 m2.
- The planet gear spacing angle may be greater than or equal to 9.0x10-1 rad. The planet gear spacing angle may be in the range between 9.0x10-1 rad to 2.1 rad.
- The radial bending stiffness of the gearbox support may be greater than or equal to 1.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 2.0x107 N/m. The radial bending stiffness of the gearbox support may be greater than or equal to 3.0x107 N/m. The radial bending stiffness of the gearbox support may be in the range from 1.0x107 N/m to 4.0x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 2.0x107 N/m to 3x108 N/m. The radial bending stiffness of the gearbox support may be in the range from 3.0x107 N/m to 2.0x108 N/m.
- The tilt stiffness of the gearbox support may be greater than or equal to 1.2x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 2.4x105 Nm/rad. The tilt stiffness of the gearbox support may be greater than or equal to 3.9x105 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 1.2x105 Nm/rad to 2.1x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 2.4x105 Nm/rad to 1.6x107 Nm/rad. The tilt stiffness of the gearbox support may be in the range from 3.9x105 Nm/rad to 9.0x106 Nm/rad.
- The gearbox may be in a star configuration.
- The maximum take-off conditions may be as defined anywhere herein, for example as operating at maximum take-off thrust for the engine at ISA sea level pressure and temperature +15°C with a fan inlet velocity of between 0.25 and 0.27 Mn, and optionally at 0.25 Mn.
- The cruise conditions may be as defined anywhere herein. The cruise conditions may mean the conditions at mid-cruise of an aircraft to which the engine is attached. The cruise conditions may be conditions experienced by the aircraft and engine at the midpoint between top of climb and start of decent.
- The forward speed of the gas turbine engine at the cruise conditions may be in the range of from Mn 0.75 to Mn 0.85. The forward speed of the gas turbine engine at the cruise conditions may be Mn 0.8.
The cruise conditions may correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11582 m and a forward Mach Number of 0.8.
The cruise conditions may correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10668 m and a forward Mach Number of 0.85.
The cruise conditions may correspond to atmospheric conditions at an altitude that is in the range of from 10500 m to 11600 m, and optionally at an altitude of 11000 m. - According to a twenty-fifth aspect there is provided a propulsor for an aircraft, comprising: a fan comprising a plurality of fan blades; a gearbox; a power unit for driving the fan via the gearbox, wherein the gearbox is arranged to receive an input from the power unit via a core shaft and output drive to a fan shaft so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the propulsor, wherein:
a flight cycle ratio of: - The propulsor of the twenty-fifth aspect may have some or all of the features described above with respect to the gas turbine engine of the twenty-fourth aspect, and may be a gas turbine engine in some embodiments.
- According to another aspect, there is provided an aircraft comprising the gas turbine engine or the propulsor of the twenty-fourth aspect or the twenty-fifth aspect mounted thereon, wherein the aircraft has a maximum take-off operating condition and a cruise condition.
- According to a twenty-sixth aspect there is provided a method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein the method comprises:
operating the gas turbine engine such that a flight cycle ratio of: - The method of the twenty-sixth aspect may be a method of operating the gas turbine engine or the propulsor of the twenty-fourth aspect or twenty-fifth aspect respectively. Any of the features, ratios and parameters introduced above in connection with the twenty-fourth aspect or twenty-fifth aspect may also therefore apply to the twenty-sixth aspect. The method of the twenty-fifth aspect may comprise operating the gas turbine engine to provide propulsion for the aircraft to which it is mounted under maximum take-off conditions. The method may further comprise operating the gas turbine engine to provide propulsion during cruise conditions. Cruise conditions and max-take off conditions are as defined elsewhere herein.
- Through consideration of a range of gearbox failure modes - including tooth root bending, tooth surface scuffing, tooth surface macro and micro pitting and gearbox vibration - and the extent to which those failure modes have an impact at maximum take-off (MTO) and cruise, the inventor has discovered that providing torsional shear stresses of the gearbox support structure within the specified range provides a low weight gearbox capable of successfully transmitting the largest loads at MTO for limited durations while providing sufficient integrity against other failure mechanisms at cruise for extended durations.
- In other aspects, value ranges for the product of the components of the flight cycle ratio may be specified instead of, or as well as, value ranges for the ratio.
- According to one such aspect, the twenty-fourth aspect introduced above may be formulated as an aspect providing a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, the gearbox being an epicyclic gearbox comprising a sun gear, a plurality of planet gears, a ring gear, and a planet carrier arranged to have the plurality of planet gears mounted thereon; and a gearbox support arranged to at least partially support the gearbox within the engine, and wherein: a product (e.g. a flight cycle product) of:
- The skilled person will appreciate that method and propulsor aspects may be formulated accordingly.
- In any of the preceding aspects, any one or more of the following may apply as applicable:
The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft - The gearbox may have a gear ratio in any range disclosed herein, for example a gear ratio in the range from 3.2 to 4.5, and optionally from 3.2 to 4.0.
The gas turbine engine may have a specific thrust in any range disclosed herein, for example a specific thrust in the range from 70 to 90 NKg-1.
The gas turbine engine may have a bypass ratio at cruise conditions in any range disclosed herein, for example in the range from 12.5 to 18, and optionally from 13 to 16. - The fan may have a fan diameter greater than 240 cm and less than or equal to 380 cm. The fan may have a fan diameter greater than 300 cm and less than or equal to 380 cm.
The diameter of the fan may be in the range from 240 cm to 280 cm. The diameter of the fan may be in the range from 330 cm to 380 cm. - The cross sectional area (CSA) of the gearbox may be defined as the area of the pitch circle of the ring gear.
- The method of any of the aspects defined above may further comprise driving the gearbox with an input torque of:
- i) greater than or equal to 10,000 Nm, and optionally of 10,000 to 50,000 Nm at cruise; and/or
- ii) greater than or equal to 28,000 Nm, and optionally of 28,000 to 135,000 Nm at max-take off conditions.
- For any parameter or ratio of parameters X claimed or disclosed herein, a limit on the values that X can take that is expressed as "X is greater than or equal to Y" can alternatively be expressed as "1/X is less than or equal to 1/Y". Any of the ratios or parameters defined in the aspects and statements above may therefore be expressed as "1/X is less than or equal to 1/Y" rather than "X is greater than or equal to Y". In such cases, zero can be taken as a lower bound.
- Various parameters of the gearbox and its mounting structure, and/or of the engine more generally, may be adjusted to allow the engine to meet the specifications of the various aspects summarised above. Comments on various such parameters are provided below.
- Regarding the stiffness (radial bending and/or tilt) of the gearbox support the inventor has found that a relatively low stiffness may be provided to isolate the gearbox from damaging loads being transmitted into it. The inventor has found that if the stiffness of the gearbox support is decreased outside of the ranges specified herein there are resulting problems with dynamic effects such as lateral vibrations - in particular, a stiffness less that is within ranges defined herein has been found to allow vibrations at low modal frequencies to be reduced or avoided (the skilled person would appreciate that the lower modal vibrations have larger amplitudes/deflections than the higher modes, and so are more important to avoid). This may be a function of the size of the gearbox and its configuration. The inventor has also found that the maximum stiffnesses provided by the ranges defined herein allows the reduction or avoidance of damaging loads being transmitted to the gearbox from the fan. This may similarly vary according to size and gearbox configuration.
- The inventor has found that decreasing the radial bending and/or tilt stiffness of the fan shaft (at the input to the fan or the output of the gearbox) outside of the ranges defined herein would lead to undesirable dynamic effects such as lateral vibration. In particular, the minimum stiffness defined by the ranged specified herein allows vibrations at low modal frequencies to be reduced or avoided (the skilled person would appreciate that the lower modal vibrations have larger amplitudes/deflections than the higher modes, and so are more important to avoid). This may be a function of the size of the gearbox and its configuration.
- The inventor has also found that an upper limit of the fan shaft radial bending and/or tilt stiffness is affected by the fundamental properties of the material or materials from which it is made. For example, a maximum stiffness is affected by the engineering limit of the material from which it is made. The materials from which the fan shaft is made (often steels) may, for example, have a Young's modulus in the range from 100 to 250 GPa, or 105 to 215 GPa, and optionally around 210 GPa - different grades of steel, or other types of metal, may be selected to achieve different stiffnesses for the same size and geometry. For example, steels with a Young's modulus in the range 190 to 215 GPa, titanium alloys with a Young's modulus in the range 105 to 120 GPa, or a metal such as titanium with a Young's modulus of around 110 GPa may be used in various embodiments. The inventor has found increasing the stiffness outside of the ranges defined herein using materials such as these would lead to excessive weight with no practical gain in performance.
- The inventor has discovered that the range of the torsional strength of the gearbox support defined herein provides a desired level of torque capacity to provide sufficient reliability, while not being so great as to add excessive weight to the engine with no practical performance gain.
- Regarding the other properties of the gearbox support such as its torsional shear stress, the inventor has found that the ranges specified herein provide the improvements in performance described above, but without leading to excessive increases in overall engine weight.
- As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.
- The gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed). The output from the gearbox may be directly to a fan shaft, or indirectly to the fan shaft, for example via a spur shaft and/or gear.
- The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft
- In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.
- The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.
- The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.
- In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).
- The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.
- The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.
Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fanblade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.
The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm. - The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.
- In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip 2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.
- Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of
form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case. - The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.
- Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-1s, 105 Nkg-1s, 100 Nkg-1s, 95 Nkg-1s, 90 Nkg-1s, 85 Nkg-1s or 80 Nkg-1s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg-1s to 100 Nkg-1s, or 85 Nkg-1s to 95 Nkg-1s. Such engines may be particularly efficient in comparison with conventional gas turbine engines.
- A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa,
temperature 30 degrees C), with the engine static. - In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K.
- The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.
- A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.
- A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.
- The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.
- The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.
- As used herein, a maximum take-off (MTO) condition has the conventional meaning. Maximum take-off conditions may be defined as operating the engine at International Standard Atmosphere (ISA) sea level pressure and temperature conditions +15°C at maximum take-off thrust at end of runway, which is typically defined at an aircraft speed of around 0.25Mn, or between around 0.24 and 0.27 Mn. Maximum take-off conditions for the engine may therefore be defined as operating the engine at a maximum take-off thrust (for example maximum throttle) for the engine at ISA sea level pressure and temperature +15°C with a fan inlet velocity of 0.25 Mn.
- As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint - in terms of time and/or distance- between top of climb and start of descent. Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.
- In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.
- Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.
- Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.
- Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m).
- In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.
- According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.
- According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).
- According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.
- Whilst in the arrangements described herein the source of drive for the propulsive fan is provided by a gas turbine engine, the skilled person will appreciate the applicability of the gearbox configurations disclosed herein to other forms of aircraft propulsor comprising alternative drive types. For example, the above-mentioned gearbox arrangements may be utilised in aircraft propulsors comprising a propulsive fan driven by an electric motor. In such circumstances, the electric motor may be configured to operate at higher rotational speeds and thus may have a lower rotor diameter and may be more power-dense. The gearbox configurations of the aforesaid aspects may be employed to reduce the rotational input speed for the fan or propeller to allow it to operate in a more favourable efficiency regime. Thus, according to an aspect, there is provided an electric propulsion unit for an aircraft, comprising an electric machine configured to drive a propulsive fan via a gearbox, the gearbox and/or its inputs/outputs/supports and/or the structure by which the fan shaft driving the fan is supported being as described and/or claimed herein.
- The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.
- As used herein, a range "from value X to value Y" or "between value X and value Y", or the like, denotes an inclusive range; including the bounding values of X and Y. As used herein, the term "axial plane" denotes a plane extending along the length of an engine, parallel to and containing an axial centreline of the engine, and the term "radial plane" denotes a plane extending perpendicular to the axial centreline of the engine, so including all radial lines at the axial position of the radial plane. Axial planes may also be referred to as longitudinal planes, as they extend along the length of the engine. A radial distance or an axial distance is therefore a distance in a radial or axial plane, respectively.
- Embodiments will now be described by way of example only, with reference to the Figures, in which:
-
Figure 1 is a sectional side view of a gas turbine engine; -
Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; -
Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; -
Figure 4 is a schematic diagram illustrating the radial bending stiffness of a cantilevered beam; -
Figure 5 is a schematic diagram illustrating the tilt stiffness of a cantilevered beam; -
Figure 6 is a schematic diagram illustrating the torsional stiffness of a shaft; -
Figure 7 is a close up sectional view of the region of a gas turbine engine around its gearbox; -
Figures 8 and 9 are schematic diagrams illustrating how the radial bending stiffness of a gearbox support may be defined; -
Figures 10 and 11 are schematic diagrams illustrating how the tilt stiffness of a gearbox support may be defined; -
Figure 12 shows a schematic sectional view of a gearbox support used with a gearbox in a star configuration; -
Figure 13 shows a schematic sectional view of a gearbox support used with a gearbox in a planetary configuration; -
Figure 14 is a schematic diagram illustrating an alternative interface between a fan shaft and fan; -
Figures 15 and 16 are schematic diagrams illustrating how fan shaft end radial bending stiffnesses may be defined; -
Figures 17 and 18 are schematic diagrams illustrating how fan shaft end tilt stiffnesses may be defined; -
Figure 19 shows an aircraft having a gas turbine engine attached to each wing; -
Figure 20 shows a method of operating a gas turbine engine on an aircraft; and -
Figure 21 shows a graph of applied load against displacement to illustrate measurement of the stiffness of a component. -
Figure 1 illustrates agas turbine engine 10 having a principalrotational axis 9. Theengine 10 comprises anair intake 12 and apropulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. Thegas turbine engine 10 comprises a core 11 that receives the core airflow A. Theengine core 11 comprises, in axial flow series, alow pressure compressor 14, a high-pressure compressor 15,combustion equipment 16, a high-pressure turbine 17, alow pressure turbine 19 and acore exhaust nozzle 20. Anacelle 21 surrounds thegas turbine engine 10 and defines abypass duct 22 and abypass exhaust nozzle 18. The bypass airflow B flows through thebypass duct 22. Thefan 23 is attached to and driven by thelow pressure turbine 19 via ashaft 26 and anepicyclic gearbox 30. - In use, the core airflow A is accelerated and compressed by the
low pressure compressor 14 and directed into thehigh pressure compressor 15 where further compression takes place. The compressed air exhausted from thehigh pressure compressor 15 is directed into thecombustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure andlow pressure turbines nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives thehigh pressure compressor 15 by a suitable interconnectingshaft 27. Thefan 23 generally provides the majority of the propulsive thrust. Theepicyclic gearbox 30 is a reduction gearbox. - An exemplary arrangement for a geared fan
gas turbine engine 10 is shown inFigure 2 . The low pressure turbine 19 (seeFigure 1 ) drives theshaft 26, which is coupled to a sun wheel, or sun gear, 28 of theepicyclic gear arrangement 30. Radially outwardly of thesun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by aplanet carrier 34. Theplanet carrier 34 constrains the planet gears 32 to precess around thesun gear 28 in synchronicity whilst enabling eachplanet gear 32 to rotate about its own axis. Theplanet carrier 34 is coupled vialinkages 36 to thefan 23 in order to drive its rotation about theengine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus orring gear 38 that is coupled, vialinkages 40, to a stationary supportingstructure 24. - The
linkages 36 may be referred to as afan shaft 36, thefan shaft 36 optionally comprising two or more shaft portions coupled together. For example, thefan shaft 36 may comprise a gearboxoutput shaft portion 36a extending from thegearbox 30 and afan portion 36b extending between the gearbox output shaft portion and thefan 23. In the embodiment shown inFigures 1 and 2 , thegearbox 30 is a planetary gearbox and the gearboxoutput shaft portion 36a is connected to the planet carrier 34 - it may therefore be referred to as acarrier output shaft 36a. Instar gearboxes 30, the gearboxoutput shaft portion 36a may be connected to the ring gear 38 - it may therefore be referred to as aring output shaft 36a. In the embodiment shown inFigures 1 and 2 , thefan portion 36b of thefan shaft 36 connects the gearboxoutput shaft portion 36a to thefan 23. The output of thegearbox 30 is therefore transferred to thefan 23, to rotate the fan, via thefan shaft 36. In alternative embodiments, thefan shaft 36 may comprise a single component, or more than two components. Unless otherwise indicated or apparent to the skilled person, anything described with respect to anengine 10 with astar gearbox 30 may equally be applied to an engine with aplanetary gearbox 30, and vice versa. - Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting
shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature is used, thefan 23 may be referred to as a first, or lowest pressure, compression stage. - The
epicyclic gearbox 30 is shown by way of example in greater detail inFigure 3 . Each of thesun gear 28, planet gears 32 andring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated inFigure 3 . There are fourplanet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetaryepicyclic gearbox 30 generally comprise at least three planet gears 32. - The
epicyclic gearbox 30 illustrated by way of example inFigures 2 and3 is of the planetary type, in that theplanet carrier 34 is coupled to an output shaft vialinkages 36, with thering gear 38 fixed. However, any other suitable type ofepicyclic gearbox 30 may be used. By way of further example, theepicyclic gearbox 30 may be a star arrangement, in which theplanet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement thefan 23 is driven by thering gear 38. By way of further alternative example, thegearbox 30 may be a differential gearbox in which thering gear 38 and theplanet carrier 34 are both allowed to rotate. - It will be appreciated that the arrangement shown in
Figures 2 and3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating thegearbox 30 in theengine 10 and/or for connecting thegearbox 30 to theengine 10. By way of further example, the connections (such as thelinkages Figure 2 example) between thegearbox 30 and other parts of the engine 10 (such as theinput shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility as defined or claimed elsewhere herein. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement ofFigure 2 . For example, where thegearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example inFigure 2 (for example as described in connection with other embodiments disclosed herein which have a star gearbox arrangement). - Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.
- Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).
- Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in
Figure 1 has asplit flow nozzle bypass duct 22 has itsown nozzle 18 that is separate to and radially outside thecore engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through thebypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. - The geometry of the
gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction inFigure 1 ), and a circumferential direction (perpendicular to the page in theFigure 1 view). The axial, radial and circumferential directions are mutually perpendicular. - The following general definitions of stiffnesses may be used herein:
- A radial bending stiffness is a measure of deformation for a given force applied in any one selected radial direction (i.e. any direction perpendicular to and passing through the engine axis). The radial bending stiffness is defined with reference to
Figure 4 in terms of the deformation of acantilevered beam 401. As illustrated inFigure 4 , a force, F, applied at the free end of the beam in a direction perpendicular to the longitudinal axis of the beam causes a linear perpendicular deformation, δ. The radial bending stiffness is the force applied for a given linear deformation i.e. F/δ. In the present application, the radial bending stiffness is taken relative to the rotational axis of theengine 9, and so relates to the resistance to linear deformation in a radial direction of the engine caused by a radial force. The beam, or equivalent cantilevered component, extends along the axis of rotation of the engine, the force, F, is applied perpendicular to the axis of rotation of the engine, along any radial direction, and the displacement δ is measured perpendicular to the axis of rotation, along the line of action of the force. The radial bending stiffness as defined herein has SI units of N/m. In the present application, unless otherwise stated, the radial bending stiffness is taken to be a free-body stiffness i.e. stiffness measured for a component in isolation in a cantilever configuration, without other components present which may affect its stiffness. - When the force is applied perpendicular to the cantilevered beam, and at the free end of the beam, the resultant curvature is not constant but rather increases towards the fixed end of the beam.
- A tilt stiffness is defined with reference to
Figure 5 , which shows the resulting deformation of acantilevered beam 401 under a moment M applied at its free end. The tilt stiffness is a measure of the resistance to rotation of a point on the component at which a moment is applied. As can be seen inFigure 5 , an applied moment at the free end of the cantilevered beam induces a constant curvature along the length of the beam between its free and fixed ends. The applied moment M causes a rotation θ of the point at which it is applied. For any component of constant section (like the beam), the angle θ is constant along the length of the component. The tilt stiffness as defined herein therefore has SI units of Nm/rad. - Torsional stiffness is a measure of deformation for a given torque.
Figure 6 illustrates the definition of the torsional stiffness of ashaft 401 or other body. A torque, τ, applied to the free end of the beam causes a rotational deformation, θ (e.g. twist) along the length of the beam. The torsional stiffness is the torque applied for a given angle of twist i.e. τ/θ. The torsional stiffness has SI units of Nm/rad. - The following general definitions of other parameters may also be used herein:
- Torque, which may also be referred to as moment, is the rotational equivalent of linear force, and can be thought of as a twist to an object. The magnitude, τ, of torque, τ , of a body depends on three quantities: the force applied (F), the lever arm vector connecting the origin to the point of force application (r), and the angle (A) between the force and lever arm vectors:
- τ is the torque vector and τ is the magnitude of the torque;
- r is the position vector or "lever arm" vector (a vector from the selected point on the body to the point where the force is applied);
- F is the force vector;
- × denotes the cross product; and
- A is the angle between the force vector and the lever arm vector (sin(A) is therefore one when the force vector is perpendicular to the position vector, such that τ = rF, i.e. magnitude of the force multiplied by distance between the selected point on the body and the point of application of the force).
- The net torque on a body determines the rate of change of the body's angular momentum.
- Moment of inertia, otherwise known as angular mass or rotational inertia, is a quantity that determines the torque needed for a desired angular acceleration of a body about a rotational axis - this is substantially equivalent to how mass determines the force needed for a particular acceleration.
- Moment of inertia depends on the body's mass distribution and the axis chosen, with larger moments requiring more torque to change the body's rotation rate. Moment of inertia has dimensions of [mass]x[distance]2 and may be expressed in units of kilogram meter squared (kg. m2). Moment of inertia I is defined as the ratio of the net angular momentum L of a body to its angular velocity ω around a principal axis:
-
- For bodies constrained to rotate in a plane, only the moment of inertia about an axis perpendicular to the plane matters, and I can therefore be represented as a scalar value. The skilled person would appreciate that a fan of a gas turbine engine (and more generally a fan rotor of the gas turbine engine comprising the fan disc and blades, and optionally also the fan shaft and/or other related components) is constrained to rotate in only one plane - a plane perpendicular to the engine axis - and that the fan's moment of inertia can therefore be defined by a single, scaler, value.
- The fan's moment of inertia about the engine axis can therefore be measured or defined using any standard methodology
- Shear stress is the component of stress coplanar with a material cross section; the stress tending to produce shear. Shear stress arises from a force vector component parallel to the cross section of the material. Shear stress may be defined as the external force acting on an object or surface parallel to the slope or plane in which it lies.
- When a shaft or other body is subjected to a torque, or twisting, a shearing stress is produced in the body - the shearing stress may be referred to as a torsional shear stress as it results from a torque. The shear stress varies from zero along the axis of the torque rotation to a maximum at the part of the body furthest from the axis; a radial distance from the axis at which the shear stress is to be measured is therefore selected. In the arrangements being described, a mid-height of the component of which the torsional shear stress is being measured is selected. For example, if the component has an outer radius of 20 cm, a mid-height position of 10 cm from the
axis 9 is selected. - Torsional shear stress has dimensions of [force]/[distance]2 and may be expressed in units of Pascals (Pa) or Newtons/metre squared (N/m2).
- Shear strength is the strength of a body against the body failing in shear - Shear strength is a material's ability to resist forces that can cause the internal structure of the material to slide against itself. A shear load is a force that tends to produce a sliding failure on a material along a plane that is parallel to the direction of the force.
- The ultimate shear strength of a material is the maximum shear stress that may be sustained before the material will rupture. The proof strength of a material is the stress at which a particular degree of permanent deformation occurs - for example, a 0.2% proof strength is the stress at which a 0.2% permanent deformation occurs.
- The strength of the body against shearing when a rotational, or twisting, shear load is applied may be referred to as torsional shear strength. The skilled person would appreciate that the shear strength of a component is important for selecting the dimensions and materials to be used for the manufacture or construction of the component. Shear strength has the same units as torsional shear stress, as it is a maximum supportable shear stress; it may therefore be expressed in units of Pa or N/m2.
- As used herein, the listed shear strength of a material is the 0.2% proof shear strength of the material unless otherwise specified. The skilled person would appreciate that this is lower than the ultimate shear strength.
- Shear strength of a material generally varies with temperature. The shear strength as used herein may be defined at room temperature.
- For metals (including metal alloys), shear strength generally decreases gradually with increasing temperature until a threshold temperature is reached, beyond which the material strength decreases rapidly. The shear strength may therefore be approximately constant below the threshold temperature. The threshold temperature may be around 400-500°C for various steel grades. As the temperatures in and around the
gearbox 30 generally do not exceed 120°C, well below the threshold temperature for likely materials for gearbox, gearbox support structure, and shaft construction, the shear strength may not be significantly different from the room temperature shear strength - the choice of a specific temperature for strength assessment may therefore have little effect. - The torsional strength of a specific component is defined as the ability of a component to withstand an applied torque without failure (i.e. without yield or structural failure). The torsional strength is therefore a measure of the torque capacity of a component. The torsional strength may be measured by applying a varying torque to a component and determining the torque at which the component fails. The torsional strength can be considered as the torque applied to the component at the point where the shear strength of that component is reached causing it to fail. The torsional strength as defined herein has dimensions of [force] x [distance] and may be expressed in units of Nm.
- More specific definitions of stiffnesses and other parameters relating to embodiments described herein are provided below for ease of understanding.
-
Figure 7 shows a region of theengine core 11 around the gearbox in close up. The same reference numbers have been used for components corresponding to those shown inFigures 1 to 3 . In the arrangement shown inFigure 7 thegearbox 30 has a star arrangement, in which thering gear 38 is coupled to thefan shaft 36 and thecarrier 34 is held in a fixed position relative to the static structure of the engine core (e.g. relative to the stationary supporting structure 24). - The
fan shaft 36 is mounted within the engine by a fanshaft mounting structure 503. The fanshaft mounting structure 503 comprises at least two bearings connected to or otherwise in engagement with the fan shaft at points spaced apart axially along the length of the engine. The fanshaft mounting structure 503 may take a number of different forms, and may comprise one or more separate supporting structures provided to support the fan shaft. It may also include other structures provided to support the fan shaft such as inter-shaft bearings. It therefore includes any supporting structure that extends between a bearing in contact with the fan shaft and a stationary structure of the engine (e.g. of the engine core). - In the arrangement shown in
Figure 7 , the fanshaft mounting structure 503 comprises two bearings, a first supportingbearing 506a and a second supportingbearing 506b, via which it is coupled to thefan shaft 36. The supportingbearings fan shaft 36. In the described arrangement, both supportingbearings gearbox 30. In other arrangements, one of the two supportingbearing fan shaft 36 may be located at a position rearward of thegearbox 30. In yet other arrangements, more than two supporting bearings may be provided as part of the fan shaft mounting structure or fan shaft mounting structure. - The
engine core 11 comprises a gearbox support 40 (corresponding to the linkage described with reference toFigure 2 ) arranged to support or mount (e.g. to at least partially support or mount) thegearbox 30 in a fixed position within the engine. The gearbox support is coupled at a first end to the stationary supportingstructure 24 which extends across thecore duct 502 carrying the core airflow A as illustrated inFigure 7 . In the presently described arrangement, thestationary support structure 24 is an engine section stator (ESS) that acts as both a structural component to provide a stationary mounting for core components such as the gearbox support, and as a guide vane provided to direct airflow from thefan 23. In other embodiments, the stationary supportingstructure 24 may comprise a strut extending across the core gas flow path and a separate stator vane provided to direct airflow. In the presently described arrangement, thegearbox support 40 is coupled at a second end to theplanet carrier 34. Thegearbox support 40 therefore acts against rotation of theplanet carrier 34 relative to the static structure of the engine core (e.g. relative to the stationary supporting structure 24). - In embodiments where the
gearbox 30 is in a planetary arrangement, thegearbox support 40 is coupled to thering gear 38 so as to resist its rotation relative to the static structure of the engine core (e.g. relative to the stationary supporting structure 24). - The
gearbox support 40 is defined between the point at which it connects to the gearbox (e.g. to the planet carrier in the presently described arrangement) and a point at which it connects to the stationary supportingstructure 24. The gearbox support may be formed by any number of separate components providing a coupling between those two points. - The
gearbox support 40 has a degree of flexibility characterized by its radial bending stiffness and its tilt stiffness. - The radial bending stiffness of the
gearbox support 40 is defined with reference toFigures 8 and 9 . The radial bending stiffness can be considered to represent the resistance of the gearbox support to a force applied to it in a radial direction of the engine. The radial bending stiffness is determined by treating thegearbox support 40 as a free body that is fixed at its point of connection with the stationary supportingstructure 24 of the engine, and which has a radial force F1 applied at its point of connection with thegearbox 30 as illustrated inFigure 8 . Deformation of thegearbox support 40 caused by the applied force F1 is illustrated inFigure 9 , with the shape of the support with no force applied shown in broken lines for comparison. The radial bending stiffness is defined in terms of the radial displacement, δ1, of the gearbox support at the position at which the force F1 is applied. The force, F1, is shown radially towards theengine axis 9 inFigure 8 , but could equivalently be a force radially away from theengine axis 9. The radial bending stiffness of thegearbox support 40 is therefore given by F1/δ1. Thegearbox 30 is shown for reference inFigure 9 remaining in a stationary position despite the deformation of thegearbox support 40 to which it is connected - this is for illustration purposes only and reflects thegearbox support 40 being treated as a free body in order to determine the stiffness. - In various embodiments, the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x107 N/m, and optionally greater than or equal to 2.0x107 N/m or greater than or equal to 3.0x107 N/m.
- In some embodiments , for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness of the gearbox support may be greater than or equal to 1.0x107 N/m or greater than or equal to 2.7x107 N/m. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness of the gearbox support may be greater than or equal to 3.2x107 N/m or greater than or equal to 4.0x107 N/m
- In various embodiments, the radial bending stiffness of the gearbox support may be in the range from 1.0x107 N/m to 4.0x108 N/m, and optionally in the range from 2.0x107 N/m to 3x108 N/m or further optionally in the range from 3.0x107 N/m to 2.0x108 N/m.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness of the gearbox support may be in the range from 1.0x107 N/m to 3.1x108 N/m and optionally in the range from 2.7x107 N/m to 3.7x107 N/m (and may be equal to 3.2x107 N/m).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness of the gearbox support may be in the range from 3.2x107 N/m to 4.0x108 N/m, and optionally in the range from 4.0x107 N/m to 5.0x107 N/m (and may be equal to 4.5x107 N/m).
- The tilt stiffness of the
gearbox support 40 is defined with reference toFigures 10 and 11 . The tilt stiffness can be considered to represent the resistance of the gearbox support to an applied moment. The tilt stiffness is determined by treating thegearbox support 40 as a free body that is fixed at its point of connection with the stationary supportingstructure 24. In order to measure the tilt stiffness, a moment M1 is applied at the point of connection between thegearbox support 40 and thegearbox 30 as illustrated inFigure 10 . Deformation of thegearbox support 40 caused by the applied moment M1 is illustrated inFigure 11 , with the shape of the support with no moment applied shown in broken lines. The tilt stiffness is defined in terms of the angular displacement, θ1, of thegearbox support 40 at the position at which the moment M1 is applied. The tilt stiffness is therefore given by M1/θ1. Thegearbox 30 is again shown for reference inFigure 11 remaining in a stationary position despite the deformation of thegearbox support 40 to which it is connected. - In various embodiments, the tilt stiffness of the gearbox support may be greater than or equal to 1.2x105 Nm/rad, and optionally greater than or equal to 2.4x105 Nm/rad or optionally greater than or equal to 3.9x105 Nm/rad.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness of the gearbox support may be greater than or equal to 1.2x105 Nm/rad or greater than or equal to 4.5x105 Nm/rad. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness of the gearbox support may be greater than or equal to 5.0x105 Nm/rad or greater than or equal to 6.0x105 Nm/rad.
- In various embodiments, the tilt stiffness of the gearbox support may be in the range from 1.2x105 Nm/rad to 2.1x107 Nm/rad, and optionally in the range from 2.4x105 Nm/rad to 1.6x107 Nm/rad, and further optionally in the range from 3.9x105 Nm/rad to 9.0x106 Nm/rad.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness of the gearbox support may be in the range from 1.2x105 Nm/rad to 7.0x106 Nm/rad and optionally in the range from 4.5x105 Nm/rad to 6.5x105 Nm/rad (and may be equal to 5.5x105 Nm/rad).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness of the gearbox support may be in the range from 5.0x105 Nm/rad to 2.1x107 Nm/rad, and optionally in the range from 6.0x105 Nm/rad to 2.6x106 Nm/rad (and may be equal to 1.6x106 Nm/rad).
- As described elsewhere herein the gearbox support is provided to locate the gearbox within the engine. In order to resist relative rotation of the gearbox compared to the stationary structure of the engine a torque is transmitted through the gearbox support. This varies over the operating cycle of the engine as different levels of torque are transmitted through the gearbox (as defined elsewhere herein) at different stages of the flight cycle of the aircraft to which the engine is mounted.
- The torque transmitted through the gearbox support is defined as the torque at the point of connection between the
gearbox support 40 and thegearbox 30. - The torsional strength of the
gearbox support 40 is defined as the level of torque applied at the point of connection between thegearbox support 40 and thegearbox 20 that would result in failure of the gearbox support. - When measuring the torsional strength or torque transmitted through the
gearbox support 40 the gearbox support is considered to be a free body that is fixed at its point of connection with the stationary supportingstructure 24 of the engine with a torque applied at the point of connection to the gearbox (i.e. in a similar manner as the application of the force or moment used to determine the radial bending stiffness and tilt stiffness of the gearbox support). - In the arrangement illustrated in
Figure 7 , the point of connection between thegearbox support 40 and thegearbox 30 is the point of connection to theplanet carrier 34 of the gearbox. Where the gearbox is in a planetary configuration, this point of connection would be between thegearbox support 40 and thering gear 38. - In various embodiments, the torsional strength of the gearbox support may be greater than or equal to 1.60x105 Nm, and optionally greater than or equal to 1.8x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torsional strength of the gearbox support may be greater than or equal to 1.8x105 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torsional strength of the gearbox support may be greater than or equal to 4.0x105 Nm or greater than or equal to 5.5x105 Nm.
- In various embodiments, the torsional strength of the gearbox support may be in the range from 1.60x105 Nm to 2.00x107 Nm, and optionally in the range from 1.8x105 Nm to 1.5x106 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torsional strength of the gearbox support may be in the range from 1.8x105 Nm to 7.0x105 Nm and optionally in the range from 1.8x105 Nm to 2.6x105 Nm (and may be equal to 2.2x105 Nm).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torsional strength of the gearbox support may be in the range from 4.0x105 Nm to 2.0x107 Nm and optionally in the range from 5.5x105 Nm to 7.5x105 Nm (and may be equal to 6.5x105 Nm).
- In various embodiments, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 6.00x104 Nm, and optionally greater than or equal to 7.2x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 7.0x104 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.8x105 Nm.
- In various embodiments, the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 6.00x104 Nm to 5.00x105 Nm, and optionally in the range from 7.2x104 Nm to 4.2x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 7.0x104 Nm to 1.9x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.8x105 Nm to 4.5x105 Nm.
- The values in the paragraphs above may apply to any gearbox configuration (i.e. star or planetary, or other gearbox arrangements).
- In various embodiments, for example in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.10x105 Nm, and optionally greater than or equal to 1.3x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.2x105 Nm or greater than or equal to 1.4x105 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 3.0x105 Nm or is greater than or equal to 3.4x105 Nm.
- In various embodiments, for example in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.10x105 Nm to 5.00x105 Nm, and optionally in the range from 1.3x105 Nm to 4.2x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.2x105 Nm to 1.9x105 Nm and optionally in the range from 1.4x105 Nm to 1.8x105 Nm (and may be equal to 1.6x105 Nm).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or the gearbox is in a star configuration, torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 3.0x105 Nm to 4.5x105 Nm and optionally in the range from 3.4x105 Nm to 4.2x105 Nm (and may be equal to 3.8x105 Nm).
- In various embodiments, for example in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 6.00x104 Nm, and optionally greater than or equal to 7.2x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 7.0x104 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be greater than or equal to 1.8x105 Nm.
- In various embodiments, for example in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 6.00x104 Nm to 3.00x105 Nm, and optionally in the range from 7.2x104 Nm to 2.6x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 7.0x104 Nm to 1.1x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or in which the gearbox is in a planetary configuration, torque transmitted through the gearbox support at maximum take-off conditions may be in the range from 1.8x105 Nm to 2.5x105 Nm.
- In various embodiments, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.00x104 Nm, and optionally greater than or equal to 2.2x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.2x104 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 6.8x104 Nm.
- In various embodiments, the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.00x104 Nm to 2.00x105 Nm, and optionally in the range from 2.2x104 Nm to 1.7x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.2x104 Nm to 6.6x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox support at cruise conditions may be in the range from 6.8x104 Nm to 1.8x105 Nm.
- The values in the paragraphs above may apply to any gearbox configuration (i.e. star or planetary, or other gearbox arrangements).
- In various embodiments, for example in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 4.00x104 Nm, and optionally greater than or equal to 4.8x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 4.5x104 Nm or greater than or equal to 5.0x104 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or the gearbox is in a star configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 1.2x105 Nm.
- In various embodiments, for example in which the gearbox is in a star configuration, the torque transmitted through the gearbox support at cruise conditions may be in the range from 4.00x104 Nm to 2.00x105 Nm, and optionally in the range from 4.8x104 Nm to 1.7x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or the gearbox is in a star configuration, the torque transmitted through the gearbox support at cruise conditions may be in the range from 4.5x104 Nm to 6.6x104 Nm and optionally in the range from 5.0x104 Nm to 6.0x104 Nm (and may be equal to 5.5x104 Nm).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or the gearbox is in a star configuration, torque transmitted through the gearbox support at cruise conditions maybe in the range from 1.2x105 Nm to 1.8x105 Nm and optionally in the range from 1.2x105 Nm to 1.8x105 Nm (and may be equal to 1.5x105 Nm).
- In various embodiments, for example in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.00x104 Nm, and optionally greater than or equal to 2.2x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 2.2x104 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at cruise conditions may be greater than or equal to 6.8x104 Nm.
- In various embodiments, for example in which the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.00x104 Nm to 1.30x105 Nm, and optionally in the range from 2.2x104 Nm to 1.1x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm and/or the gearbox is in a planetary configuration, the torque transmitted through the gearbox support at cruise conditions may be in the range from 2.2x104 Nm to 3.3x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm and/or the gearbox is in a planetary configuration, torque transmitted through the gearbox support at cruise conditions may be in the range from 6.8x104 Nm to 1.1x105 Nm.
- The
gearbox support 40 has a torsional shear stress that also represents the resistance of thegearbox support 40 to a torque applied by thegearbox 30. The torsional shear stress is as defined elsewhere herein. - In various embodiments, the torsional shear stress of the gearbox support at maximum take-off conditions may be less than or equal to 4.90x108 N/m2, and optionally less than or equal to 3.5x108 N/m2.
- In various embodiments, the torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 1.40x108 N/m2 to 4.90x108 N/m2, and optionally in the range from 2.0x108 N/m2 to 3.5x108 N/m2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm or in the range from 330 to 380 cm, the torsional shear stress of the gearbox support at maximum take-off conditions may be in the range from 2.3x108 N/m2 to 3.7x108 N/m2 (and may be equal to 2.5x108 N/m2).
- In various embodiments, the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 7.00x107 N/m2, and optionally greater than or equal to 8.2x107 N/m2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 8.0x107 N/m2. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torsional shear stress of the gearbox support at cruise conditions may be greater than or equal to 8.5x107 N/m2 or greater than or equal to 9.0x107 N/m2.
- In various embodiments, the torsional shear stress of the gearbox support at cruise conditions may be in the range from 7.00x107 N/m2 to 1.90x108 N/m2, and optionally in the range from 8.2x107 N/m2 to 1.5x108 N/m2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torsional shear stress of the gearbox support at cruise conditions may be in the range from 8.0x107 N/m2 to 1.5x108 N/m2 and optionally in the range from 8.0x107 N/m2 to 9.2x107 N/m2 (and may be equal to 8.6x107 N/m2).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torsional shear stress of the gearbox support at cruise conditions may be in the range from 8.5x107 N/m2 to 1.9x108 N/m2 and optionally in the range from 9.0x107 Nm2 to 1.2x108 N/m2 (and may be equal to 9.6x107 N/m2).
- The maximum take-off conditions and cruise conditions referred to in this section may be as defined elsewhere herein.
- The shear stress may be measured through the highest loaded plane section of the component (i.e. the gearbox support). It excludes the effects of stress concentrations at small radii and holes. The skilled person will understand that for the shear stress of the gearbox support defined herein since the predominant load is a torque, the effective radius also affects the measurement of the shear as well as the area.
- The
fan 23 has a moment of inertial I F. The moment of inertia of the fan is measured based on the total mass of the rotor forming the fan, i.e. including the total mass of the plurality of fan blades, the fan hub and any support arm or other linkages provided to connect the fan to the fan shaft. The moment of inertial therefore includes all rotating components apart from the fan shaft. The moment of inertia is the mass moment of inertia or rotational inertia of the fan with respect to rotation around the principalrotational axis 9 of the engine. Rotation of the fan will cause a gyroscopic effect meaning that the fan shaft will tend to maintain a steady direction of its axis of rotation. During manoeuvring of the aircraft to which the gas turbine engine is mounted the orientation of the axis of rotation of the fan shaft will however change. The gyroscopic effect will result in a reaction force at the fan shaft mounting structure to resist the tendency of the fan shaft to maintain its orientation. The moment of inertia of the fan will have an effect on the magnitude of the gyroscopic effect produced, and so has an impact on the design of the fan shaft and the fan shaft mounting structure as is discussed elsewhere herein. - In various embodiments, the moment of inertia of the fan may be greater than or equal to 7.40x107 kgm2, and optionally greater than or equal to 8.3x107 kgm2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the moment of inertia of the fan may be greater than or equal to 7.4x107 kgm2 or 8.6x107 kgm2. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the moment of inertia of the fan may be greater than or equal to 3.0x108 kgm2 or 4.0x108 kgm2.
- In various embodiments, the moment of inertia of the fan may be in the range from 7.40x107 kgm2 to 9.00x108 kgm2, and optionally in the range from 8.3x107 kgm2 to 6.5x108 kgm2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the moment of inertia of the fan may be in the range from 7.4x107 kgm2 to 1.5x108 kgm2 and optionally in the range from 8.6x107 kgm2 to 9.6x107 kgm2 (and may be equal to 9.1x107 kgm2).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the moment of inertia of the fan may be in the range from 3.0x108 kgm2 to 9.0x108 kgm2 and optionally in the range from 4.0x108 kgm2 to 5.0x108 kgm2 (and may be equal to 4.5x108 kgm2).
- Referring again to
Figure 7 , a fan-gearboxaxial distance 110 is defined as the axial distance between the axial position P of the gearbox output position (also labelled X inFigures 15 to 18 ) and the fan axial centreline Q. The fan axial centreline is defined as the axial midpoint of the fan blades forming the fan. The gearbox output position is defined as the point of connection between thefan shaft 36 and the gearbox. This may be defined differently for different types of gearbox as is described below. - In various embodiments, the fan-gearbox axial distance may be greater than or equal to 0.35 m, and optionally greater than or equal to 0.37 m.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan-gearbox axial distance may be greater than or equal to 0.38 m or greater than or equal to 0.40 m. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan-gearbox axial distance may be greater than or equal to 0.48 m or greater than or equal to 0.50 m.
- In various embodiments, the fan-gearbox axial distance may be in the range from 0.35 m to 0.8 m, and optionally in the range from 0.37 m to 0.75 m.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan-gearbox axial distance may be in the range from 0.38 m to 0.65 m and optionally in the range from 0.40 m to 0.44 m (and may be equal to 0.42 m).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan-gearbox axial distance may be in the range from 0.48 m to 0.8 m and optionally in the range from 0.50 m to 0.66 m (and may be equal to 0.58 m).
- The
fan shaft 36 is defined as the torque transfer component that extends from the output of the gearbox to the fan input. It therefore includes part or all of any gearbox output shaft and fan input shaft that may be provided between those points. For the purposes of defining the stiffness of thefan shaft 36 it is considered to extend between a fan input position and a gearbox output position, and includes all of the torque transfer components between those points. It does not therefore include any components of the gearbox (e.g. the planet carrier or connecting plate coupled to it) which transmit discrete forces, rather than the fan shaft torque. The gearbox output position therefore may be defined as the point of connection between thefan shaft 36 and thegearbox 30. The fan input position may be defined as the point of connection between thefan shaft 36 and the fan. - Referring to
Figure 12 , where the gearbox is in a star configuration, the gearbox output position is defined as the point ofconnection 702 between thering gear 38 and thefan shaft 36. More specifically, it is the point of connection to the annulus of the ring gear 38 (with any connection component extending from the outer surface of the annulus being considered to be part of the ring gear). Where the point of connection is formed by an interface extending in a direction having an axial component, the point of connection is considered to be the axial centreline of that interface as illustrated inFigure 7 . - The
fan shaft 36 includes all torque transmitting components up to the point ofconnection 702 with thering gear 38. It therefore includes any flexible portions orlinkages 704 of thefan shaft 36 that may be provided, and any connection(s) 706 (e.g. spline connections) between them. - Where the
gearbox 30 is in a planetary configuration, the gearbox output position is again defined as the point of connection between thefan shaft 36 and thegearbox 30. An example of this is illustrated inFigure 13 , which shows a carrier comprising aforward plate 34a andrearward plate 34b, with a plurality ofpins 33 extending between them and on which the planet gears are mounted. Thefan shaft 36 is connected to theforward plate 34a via aspline connection 708. In an arrangement such as this, the gearbox output position is taken as any point on the interface between thefan shaft 36 and theforward plate 34a. Theforward plate 34a is considered to transmit discrete forces, rather than a single torque, and so is taken to be part of thegearbox 30 rather than the fan shaft. -
Figure 13 shows only one example of a type of connection between the fan shaft andplanet carrier 34. In embodiments having different connection arrangements, the gearbox output position is still taken to be at the interface between components transmitting a torque (i.e. that are part of the fan shaft) and those transmitting discrete forces (e.g. that are part of the gearbox). Thespline connection 708 is only one example of a connection that may be formed between the fan shaft and gearbox (i.e. between the fan shaft and theforward plate 34b in the presently described arrangement). In other embodiments, the interface which forms the gearbox output position may be formed by, for example, a curvic connection, a bolted joint or other toothed or mechanically dogged arrangement. - The fan input position is defined as a point on the fan shaft at the axial midpoint of the interface between the fan and the fan shaft. In the presently described arrangement, the
fan 23 comprises asupport arm 23a (as can be seen, for example, inFigure 7 ) arranged to connect thefan 23 to thefan shaft 36. Thesupport arm 23a is connected to the fan shaft by aspline coupling 36c that extends along the length of a portion of thefan shaft 36. The fan input position is defined as the axial midpoint of the spline coupling. The spline coupling shown is only one example of a coupling that may form the interface between the fan and fan shaft. In other embodiments, for example, a curvic connection, a bolted joint or other toothed or mechanically dogged arrangement may be used. For example, a flange coupling may be provided between thesupport arm 23a and thefan shaft 36. In such an embodiment, the support arm can be connected at the rear of the fan hub. In this embodiment the fan input position is the axial midpoint of the flange coupling.Figure 14 illustrates an arrangement in which an alternative coupling is provided between thefan 23 and thefan shaft 36. Similarly toFigure 7 , thefan 23 is coupled to thefan shaft 36 via asupport arm 23a. In this arrangement however, aflange coupling 36d is provided between thesupport arm 23a and thefan shaft 36. In this embodiment, thesupport arm 23a is connected at the rear of the fan hub. Theflange coupling 36d may be a curvic coupling. In other embodiments, other forms of flange coupling may be provided. In the embodiment ofFigure 14 , the fan input position is the axial midpoint (labelled Y) of the flange coupling. - The
fan shaft 36 has a degree of flexibility characterized by its radial bending stiffness and tilt stiffness. - The stiffness of the fan shaft where it couples to the
gearbox 30 is defined with reference toFigures 15 to 18 . - The radial bending stiffness of the
fan shaft 36 at the output of thegearbox 30 is measured by applying a force F2 to the fan shaft at the gearbox output position defined above (illustrated inFigure 15 ). Thefan shaft 36 is treated as a free body, and is held fixed at the position of all of bearing positions at which it is supported i.e. the first andsecond bearings Figure 15 . As a result of force F2 thefan shaft 36 deforms so that the gearbox output position is displaced by a distance of δ2 (as illustrated inFigure 16 ). The radial bending stiffness of thefan shaft 36 at the output of the gearbox is then given by F2/δ2. - In various embodiments, the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 4.00x106 N/m, and optionally greater than or equal to 3.7x107 N/m.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 3.7x107 N/m. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 3.9x107 N/m or greater than or equal to 5.0x107 N/m.
- In various embodiments, the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 4.00x106 N/m to 1.5x109 N/m, and optionally in the range from 3.7x107 N/m to 1.0x109 N/m.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 3.7x107 N/m to 5.0x108 N/m and optionally in the range from 3.7x107 N/m to 4.3x107 N/m (and may be equal to 4.0x107 N/m).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness of the fan shaft at the output of the gearbox may be in the range from 3.9x107 N/m to 1.5x109 N/m, and optionally in the range from 5.0x107 N/m to 9.0x107 N/m (and may be equal to 7.0x107 N/m).
- The tilt stiffness of the
fan shaft 36 at the output of thegearbox 30 is measured by applying a moment M2 to the fan shaft at the gearbox output position defined above (illustrated inFigure 17 ). Thefan shaft 36 is again treated as a free body, and is held fixed at the position of all of the bearing positions at which it is supported i.e. the first andsecond bearings Figure 15 . As a result of moment M2 thefan shaft 36 deforms so that the gearbox output position is displaced by an angular displacement of θ2 as illustrated inFigure 17 . The tilt stiffness of thefan shaft 36 at the output of the gearbox is then given by M2/θ2. - In various embodiments, the tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 7.00x104 Nm/rad, and optionally greater than or equal to 9.5x105 Nm/rad.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness of the fan shaft at the output of the gearbox may be greater than or equal to 9.5x105, and optionally may be greater than or equal to 9.5x105 Nm/rad. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness of the fan shaft output of the gearbox may be greater than or equal to 1.1x106 Nm/rad and optionally may be greater than or equal to 2.0x106 Nm/rad .
- In various embodiments, the tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 7.00x104 Nm/rad to 7.00x107 Nm/rad, and optionally in the range from 9.5x105 Nm/rad to 3.5x107 Nm/rad.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness of the output of the gearbox may be in the range from 9.5x105 Nm/rad to 2.0x107 Nm/rad and optionally in the range from 9.5x105 Nm/rad to 2.4x106 Nm/rad (and may be equal to 1.2x106 Nm/rad).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness of the fan shaft at the output of the gearbox may be in the range from 1.1x106 Nm/rad to 7.0x107 Nm/rad, and optionally in the range from 2.0x106 Nm/rad to 5.2x106 Nm/rad (and may be equal to 3.6x106 Nm/rad).
- The gearbox provides a torque conversion between the toque at its input (i.e. the core shaft) and its output (i.e. the fan shaft). The torque transmitted through the gearbox is defined as the torque at the output position of the gearbox (the output position being as defined elsewhere herein). The torque transmitted through the gearbox varies over the operating cycle of the engine.
- In various embodiments, the torque transmitted through the gearbox at maximum take-off conditions may be greater than or equal to 7.00x104 Nm, and optionally greater than or equal to 1.0x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox at maximum take-off conditions may be greater than or equal to 1.1x105 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox at maximum take-off conditions may be greater than or equal to 1.5x105 Nm or greater than or equal to 2.0x105 Nm.
- In various embodiments, the torque transmitted through the gearbox at maximum take-off conditions may be in the range from 7.00x104 Nm to 5.00x105 Nm, and optionally in the range from 1.0x105 Nm to 3.5x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox at maximum take-off conditions may be in the range from 1.1x105 Nm to 1.5x105 Nm and optionally in the range from 1.1x105 Nm to 1.3x105 Nm (and may be equal to 1.2x105 Nm).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox at maximum take-off conditions may be in the range from 1.5x105 Nm to 5.0x105 Nm and optionally in the range from 2.0x105 Nm to 3.8x105 Nm (and may be equal to 2.9x105 Nm).
- In various embodiments, the torque transmitted through the gearbox at cruise conditions may be greater than or equal to 2.30x104 Nm, and optionally greater than or equal to 3.1x104 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox at cruise conditions may be greater than or equal to 3.2x104 Nm or greater than or equal to 3.8x104 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque transmitted through the gearbox at cruise conditions may be greater than or equal to 7.3x104 Nm or greater than or equal to 9.8x104 Nm.
- In various embodiments, the torque transmitted through the gearbox at cruise conditions may be in the range from 2.30x104 Nm to 1.80x105 Nm, and optionally in the range from 3.1x104 Nm to 1.5x105 Nm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque transmitted through the gearbox at cruise conditions may be in the range from 3.2x104 Nm to 7.2x104 Nm and optionally in the range from 3.8x104 Nm to 4.6x104 Nm (and may be equal to 4.2x104 Nm).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, torque transmitted through the gearbox at cruise conditions may be in the range from 7.3x104 Nm to 1.8x105 Nm and optionally in the range from 9.8x104 Nm to 1.4x105 Nm (and may be equal to 1.1x105 Nm).
- The cross sectional area (CSA) of the gearbox is defined as the area of the pitch circle of the ring gear. The pitch circle of a gear is an imaginary circle that rolls without slipping with the pitch circle of any other gear with which the first gear is meshed. The pitch circle passes through the points where the teeth of two gears meet as the meshed gears rotate - the pitch circle of a gear generally passes through a mid-point of the length of the teeth of the gear. The CSA of the gearbox can be found by measuring the pitch circle diameter (PCD) of the gear. The PCD can be roughly estimated by taking the average of the diameter between tips of the gear teeth and the diameter between bases of the gear teeth.
- In various embodiments, the CSA of the gearbox may be greater than or equal to 2.4x10-1 m2 and optionally greater than or equal to 2.6x10-1 m2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the CSA of the gearbox may be greater than or equal to 2.4x10-1 m2 or greater than or equal to 2.5x10-1 m2. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the CSA of the gearbox may be greater than or equal to 4.5x10-1 m2 or greater than or equal to 5.5x10-1 m2.
- In various embodiments, the CSA of the gearbox may be in the range from 2.4x10-1 m2 to 1.10 m2, and optionally in the range from 2.6x10-1 m2 to 9.0x10-1 m2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the CSA of the gearbox may be in the range from 2.4x10-1 m2 to 5.0x10-1 m2 and optionally in the range from 2.5x10-1 m2 to 3.4x10-1 m2 (and may be equal to 2.9x10-1 m2).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the CSA of the gearbox may be in the range from 4.5x10-1 m2 to 1.1 m2, and optionally in the range from 5.5x10-1 m2 to 6.4x10-1 m2 (and may be equal to 5.9x10-1 m2).
- A planet gear spacing angle in radians (β) is defined as 2π/N, where N is the number of planet gears 32 provided in the gearbox. The planet gear spacing angle is illustrated in
Figure 3 . The planet gearbox spacing angle corresponds to the average angle (in rad) between all the adjacent pairs of planet gears. - In various embodiments, the planet gear spacing angle (β) may be greater than or equal to 9.0x10-1 rad, and optionally in the range between 9.0x10-1 rad to 2.1 rad, and further optionally in the range between 1.1 rad and 1.3 rad (and may be equal to 1.26 rad)
- The inventor has discovered that the ratios (and/or products) of some properties have a considerable impact on the operation of the gearbox and its inputs/outputs/support structure. Some or all of the below may apply to any embodiment:
In various embodiments, a radial bending stiffness to moment of inertia ratio may be defined as: - In various embodiments, the radial bending stiffness to moment of inertia ratio may be greater than or equal to 2.5x10-2 Nkg-1m-3 (i.e. (N/m)/(kg.m2)) and optionally greater than or equal to 0.05 Nkg-1m-3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg-1m-3. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.025 Nkg-1m-3.
- In various embodiments, the radial bending stiffness to moment of inertia ratio may be in the range from 2.5x10-2 Nkg-1m-3 to 6.0 Nkg-1m-3, and optionally in the range from 0.05 Nkg-1m-3 to 3.0 Nkg-1m-3, and further optionally in the range from 0.05 Nkg-1m-3 to 0.6 Nkg-1m-3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg-1m-3 to 4.0 Nkg-1m-3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness to moment of inertia ratio may be in the range from 0.025 Nkg-1m-3 to 2.0 Nkg-1m-3.
-
- In various embodiments, the radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x1014 Nkgm (i.e. (N/m).(kg.m2)), and optionally greater than or equal to 4.0x1014 Nkgm, and further optionally greater than or equal to 2.0x1015 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness to moment of inertia product may be greater than or equal to 1.5x1015 Nkgm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x1015 Nkgm.
- In various embodiments, the radial bending stiffness to moment of inertia product may be in the range from 2.0x1014 Nkgm to 1.4x1018 Nkgm, and optionally in the range from 4.0x1014 Nkgm to 7.0x1017 Nkgm, and further optionally in the range from 2.0x1015 Nkgm to 7.0x1017 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the radial bending stiffness to moment of inertia product may be in the range from 1.5x1015 Nkgm to 1.3x1017 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the radial bending stiffness to moment of inertia product may be in the range from 2.0x1015 Nkgm to 1.4x1018Nkgm.
-
- In various embodiments, the fan shaft radial bending stiffness to moment of inertia ratio may be greater than or equal to 2.5x10-2 Nkg-1m-3 (i.e. (N/m)/(kg.m2)), and optionally greater than or equal to 0.05 (N/m)/(kg.m2).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.05 Nkg-1m-3 or greater than or equal to 0.4 Nkg-1m-3. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.025 Nkg-1m-3 or optionally greater than or equal to 0.06 Nkg-1m-3.
- In various embodiments, the fan shaft radial bending stiffness to moment of inertia ratio may be in the range from 2.5x10-2 Nkg-1m-3 to 6.0 Nkg-1m-3, and optionally in the range from 0.05 Nkg-1m-3 to 3.0 Nkg-1m-3, and further optionally in the range from 0.05 Nkg-1m-3 to 0.6 Nkg-1m-3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft radial bending stiffness to moment of inertia ratio may be in the range from 0.05 Nkg-1m-3 to 0.6 Nkg-1m-3 and optionally in the range from 0.4 Nkg-1m-3 to 0.5 Nkg-1m-3 (and may be equal to 0.44 Nkg-1m-3).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft radial bending stiffness to moment of inertia ratio may be in the range from 0.025 Nkg-1m-3 to 0.6 Nkg-1m-3, and optionally in the range from 0.06 Nkg-1m-3 to 0.26 Nkg-1m-3 (and may be equal to 0.16 Nkg-1m-3).
- In various embodiments, in addition to or alternatively to the fan shaft radial bending stiffness to moment of inertia ratio, a product of the parameters making up that ratio may be defined. This product (referred to as the fan shaft radial bending stiffness to moment of inertia product) may be defined as:
- In various embodiments, the fan shaft radial bending stiffness to moment of inertia product may be greater than or equal to 3.0x1014 Nkgm (i.e. (N/m).(kg.m2)), and optionally greater than or equal to 6.0x1014 Nkgm, and further optionally greater than or equal to 2.0x1015 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x1015 Nkgm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft radial bending stiffness to moment of inertia product may be greater than or equal to 2.3x1015 Nkgm.
- In various embodiments, the fan shaft radial bending stiffness to moment of inertia product may be in the range from 3.0x1014 Nkgm to 1.4x1018 Nkgm, and optionally in the range from 6.0x1014 Nkgm to 7.0x1017 Nkgm, and further optionally in the range from 2.0x1015 Nkgm to 70x1017 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft radial bending stiffness to moment of inertia product may be in the range from 2.0x1015 Nkgm to 7.5x1016 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft radial bending stiffness to moment of inertia product may be in the range from 2.3x1015 Nkgm to 1.4x1018 Nkgm.
-
- In various embodiments, the gearbox support radial bending stiffness to moment of inertia ratio may be greater than or equal to 3.0x10-2 Nkg-1m-3 (i.e. (N/m)/(kg.m2)), and optionally greater than or equal to 0.06 Nkg-1m-3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.06 Nkg-1m-3 or greater than or equal to 0.25 Nkg-1m-3. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support radial bending stiffness to moment of inertia ratio may be greater than or equal to 0.03 Nkg-1m-3 or greater than or equal to 0.06 Nkg-1m-3.
- In various embodiments, the gearbox support radial bending stiffness to moment of inertia ratio may be in the range from 3.0x10-2 Nkg-1m-3 to 4.0 Nkg-1m-3, and optionally in the range from 0.06 Nkg-1m-3 to 2.0 Nkg-1m-3, and further optionally in the range from 0.06 Nkg-1m-3 to 0.48 Nkg-1m-3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support radial bending stiffness to moment of inertia ratio may be in the range from 0.06 Nkg-1m-3 to 4.0 Nkg-1m-3 and optionally in the range from 0.25 Nkg-1m-3 to 0.45 Nkg-1m-3 (and may be equal to 0.35 Nkg-1m-3).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support radial bending stiffness to moment of inertia ratio may be in the range from 0.03 Nkg-1m-3 to 2.0 Nkg-1m-3, and optionally in the range from 0.06 Nkg-1m-3 to 0.6 Nkg-1m-3 (and may be equal to 0.1) Nkg-1m-3.
- In various embodiments, in addition to or alternatively to the gearbox support radial bending stiffness to moment of inertia ratio, a product of the parameters making up that ratio may be defined. This product (referred to as the gearbox support radial bending stiffness to moment of inertia product) may be defined as: the radial bending stiffness of the gearbox support (40) × the moment of inertia of the fan (23)
- In various embodiments, the gearbox support radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x1014 Nkgm (i.e. (N/m).(kg.m2)), and optionally greater than or equal to 4.0x1014 Nkgm or further optionally greater than or equal to 2.0x1015 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support radial bending stiffness to moment of inertia product may be greater than or equal to 1.5x1015 Nkgm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support radial bending stiffness to moment of inertia product may be greater than or equal to 2.0x1015 Nkgm.
- In various embodiments, the gearbox support radial bending stiffness to moment of inertia product may be in the range from 2.0x1014 Nkgm to 3.0x1017 Nkgm, and optionally in the range from 4.0x1014 Nkgm to 1.3x1017 Nkgm, and yet further optionally in the range from 2.0x1015 Nkgm to 1.3x1017 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support radial bending stiffness to moment of inertia product may be in the range from 1.5x1015 Nkgm to 1.3x1017 Nkgm.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support radial bending stiffness to moment of inertia product may be in the range from 2.0x1015 Nkgm to 3.0x1017 Nkgm.
-
- In various embodiments, the tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10-4 Nrad-1kg-1m-1 (i.e. (Nm/rad)/(kg.m2)), and optionally greater than or equal to 1.0x10-3 Nrad-1kg-1m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10-1 Nrad-1kg-1m-1. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10-4 Nrad-1kg-1m-1.
- In various embodiments, the tilt stiffness to moment of inertia ratio may be in the range from 4.0x10-4 Nrad-1kg-1m-1 to 2.7x10-1 Nrad-1kg-1m-1, and optionally in the range from 1.0x10-3 Nrad-1kg-1m-1 to 0.1 Nrad-1kg-1m-1, and further optionally in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.5x10-2 Nrad-1kg-1m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness to moment of inertia ratio may be in the range from 1.0x10-1 Nrad-1kg-1m-1 to 1.45x10-2 Nrad-1kg-1m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness to moment of inertia ratio may be in the range from 4.0x10-4 Nrad-1kg-1m-1 to 3.0x10-2 Nrad-1kg-1m-1.
-
- In various embodiments, the tilt stiffness to moment of inertia product may be greater than or equal to 3.0x1012 Nm3rad-1kg (i.e. (Nm/rad).(kg.m2)), and optionally greater than or equal to 6.0x1012 Nm3rad-1kg, and further optionally greater than or equal to 2.5x1013 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness to moment of inertia product may be greater than or equal to 2.0x1013 Nm3rad-1kg. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness to moment of inertia product may be greater than or equal to 2.5x1013 Nm3rad-1kg.
- In various embodiments, the tilt stiffness to moment of inertia product may be in the range from 3.0x1012 Nm3rad-1kg to 6.0x1016 Nm3rad-1kg, and optionally in the range from 6.0x1012 Nm3rad-1kg to 3.0x1016 Nm3rad-1kg, and further optionally in the range from 2.5x1013 Nm3rad-1kg to 3.0x1016 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the tilt stiffness to moment of inertia product may be in the range from 2.0x1013 Nm3rad-1kg to 4.0x1015 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the tilt stiffness to moment of inertia product may be in the range from 2.5x1013 Nm3rad-1kg to 6.0x1016 Nm3rad-1kg.
-
- In various embodiments, the fan shaft tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10-4 Nrad-1kg-1m-1 (i.e. (Nm/rad)/(kg.m2)), and optionally greater than or equal to 1.0x10-3 Nrad-1kg-1m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10-3 Nrad-1kg-1m-1 or greater than or equal to 3.0x10-3 Nrad-1kg-1m-1. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft tilt stiffness to moment of inertia ratio may be greater than or equal to 4.0x10-4 Nrad-1kg-1m-1 or greater than or equal to 0.7x10-3 Nrad-1kg-1m-1.
- In various embodiments, the fan shaft tilt stiffness to moment of inertia ratio may be in the range from 4.0x10-4 Nrad-1kg-1m-1 to 0.27 Nrad-1kg-1m-1, and optionally in the range from 1.0x10-3 Nrad-1kg-1m-1 to 0.1 Nrad-1kg-1m-1, and further optionally in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.5x10-2 Nrad-1kg-1m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft tilt stiffness to moment of inertia ratio may be in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.45x10-2 Nrad-1kg-1m-1 and optionally in the range from 3.0x10-3 Nrad-1kg-1m-1 to 2.3x10-2 Nrad-1kg-1m-1 (and may be equal to 1.3x10-2 Nrad-1kg-1m-1).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft tilt stiffness to moment of inertia ratio may be in the range from 4.0x10-4 Nrad-1kg-1m-1 to 1.4x10-2 Nrad-1kg-1m-1, and optionally in the range from 7.0x10-3 Nrad-1kg-1m-1 to 9.0x10-3 Nrad-1kg-1m-1 (and may be equal to 8.0x10-3 Nrad-1kg-1m-1).
-
- In various embodiments, the fan shaft tilt stiffness to moment of inertia product may be greater than or equal to 5.0x1012 Nm3rad-1kg (i.e. (Nm/rad).(kg.m2)), and optionally greater than or equal to 1.0x1013 Nm3rad-1kg or greater than or equal to 6.0x1013 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft tilt stiffness to moment of inertia product may be greater than or equal to 6.0x1013 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft tilt stiffness to moment of inertia product may be greater than or equal to 1.0x1014 Nm3rad-1kg.
- In various embodiments, the fan shaft tilt stiffness to moment of inertia product may be in the range from 5.0x1012 Nm3rad-1kg to 6.0x1016 Nm3rad-1kg, and optionally in the range from 1.0x1013 Nm3rad-1kg to 3.0x1016 Nm3rad-1kg, and further optionally in the range from 6.0x1013 Nm3rad-1kg to 3.0x1016 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan shaft tilt stiffness to moment of inertia product may be in the range from 6.0x1013 Nm3rad-1kg to 3.0x1015 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan shaft tilt stiffness to moment of inertia product may be in the range from 1.0x1014 Nm3rad-1kg to 6.0x1016 Nm3rad-1kg.
-
- In various embodiments, the gearbox support tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10-3 Nrad-1kg-1m-1 (i.e. (Nm/rad)/(kg.m2)), and optionally greater than or equal to 2.0x10-3 Nrad-1kg-1m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support tilt stiffness to moment of inertia ratio may be greater than or equal to 2.0x10-3 Nrad-l1kg-1m-1 or greater than or equal to 5.0x10-3 Nrad-1kg-1m-1. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support tilt stiffness to moment of inertia ratio may be greater than or equal to 1.0x10-3 Nrad-1kg-1m-1 or greater than or equal to 2.6x10-3 Nrad-1kg-1m-1.
- In various embodiments, the gearbox support tilt stiffness to moment of inertia ratio may be in the range from 1.0x10-3 Nrad-1kg-1m-1 to 7.0x10-2 Nrad-1kg-1m-1. and optionally in the range from 2.0x10-3 Nrad-1kg-1m-1 to 3.0x10-2 Nrad-1kg-1m-1, and further optionally in the range from 2.0x10-3 Nrad-1kg-1m-1 to 7.0x10-3 Nrad-1kg-1m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support tilt stiffness to moment of inertia ratio may be in the range from 2.0x10-3 Nrad-1kg-1m-1 to 7.2x10-3 Nrad-1kg-1m-1 and optionally in the range from 5.0x10-3 Nrad-1kg-1m-1 to 7.0x10-3 Nrad-1kg-1m-1 (and may be equal to 6.0x10-3 Nrad-1kg-1m-1).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support tilt stiffness to moment of inertia ratio may be in the range from 1.0x10-3 Nrad-1kg-1m-1 to 3.0x10-2 Nrad-1kg-1m-1, and optionally in the range from 2.6x10-3 Nrad-1kg-1m-1 to 4.6x10-3 Nrad-1kg-1m-1 (and may be equal to 3.6x10-3 Nrad-1kg-1m-1).
-
- In various embodiments, the gearbox support tilt stiffness to moment of inertia product may be greater than or equal to 3.0x1012 Nm3rad-1kg (i.e. (Nm/rad).(kg.m2)), and optionally greater than or equal to 6.0x1012 Nm3rad-1kg, and further optionally greater than or equal to 2.5x1013 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support tilt stiffness to moment of inertia product may be greater than or equal to 2.0x1013 Nm3rad-1kg. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support tilt stiffness to moment of inertia product may be greater than or equal to 2.5x1013 Nm3rad-1kg.
- In various embodiments, the gearbox support tilt stiffness to moment of inertia product may be in the range from 3.0x1012 Nm3rad-1kg to 9.0x1015 Nm3rad-1kg, and optionally in the range from 6.0x1012 Nm3rad-1kg to 4.0x1015 Nm3rad-1kg, and further optionally in the range from 2.5x1013 Nm3rad-1kg to 4.0x1015 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the gearbox support tilt stiffness to moment of inertia product may be in the range from 2.0x1013 Nm3rad-1kg to 4.0x1015 Nm3rad-1kg.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the gearbox support tilt stiffness to moment of inertia product may be in the range from 2.5x1013 Nm3rad-1kg to 9.0x1015 Nm3rad-1kg.
- In various embodiments, a product of the fan-gearbox axial distance multiplied by the moment of inertia of the fan may be defined.
- In various embodiments, the product of the fan-gearbox axial distance and the moment of inertia of the fan may be greater than or equal to 1.9x107 kgm3, and optionally greater than or equal to 2.9x107 kgm3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the product of the fan-gearbox axial distance and the moment of inertia of the fan may be greater than or equal to 2.0x107 kgm3. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the product of the fan-gearbox axial distance and the moment of inertia of the fan may be greater than or equal to 1.2x108 kgm3.
- In various embodiments, the product of the fan-gearbox axial distance and the moment of inertia of the fan may be in the range from 1.9x107 kgm3 to 6.2x108 kgm3, and optionally in the range from 2.9x107 kgm3 to 3.9x108 kgm3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the product of the fan-gearbox axial distance and the moment of inertia of the fan may be in the range from 2.0x107 kgm3 to 8.0x107kgm3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the product of the fan-gearbox axial distance and the moment of inertia of the fan may be in the range from 1.2x108 kgm3 to 6.2x108 kgm3.
- In various embodiments, a ratio given by the fan-gearbox axial distance divided by the moment of inertia of the fan may also be defined.
- In various embodiments, the fan-gearbox axial distance divided by the moment of inertia of the fan may be less than or equal to 8.8x10-9 m/kgm2 (i.e. kg-1m-1), and optionally less than or equal to 6.2x10-9 m/kgm2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan-gearbox axial distance divided by the moment of inertia of the fan may be less than or equal to 6.5x10-9 m/kgm2 or less than or equal to 5.2x10-9 m/kgm2. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan-gearbox axial distance divided by the moment of inertia of the fan may be less than or equal to 2.8x10-9 m/kgm2 or less than or equal to 1.8x10-9 m/kgm2.
- In various embodiments, the fan-gearbox axial distance divided by the moment of inertia of the fan may be in the range from 5.3x10-10 m/kgm2 to 8.8x10-9 m/kgm2, and optionally in the range from 8.8x10-10 m/kgm2 to 6.2x10-9 m/kgm2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the fan-gearbox axial distance divided by the moment of inertia of the fan may be in the range from 2.1x10-9 m/kgm2 to 6.5x10-9 m/kgm2 and optionally in the range from 4.0x10-9 m/kgm2 to 5.2x10-9 m/kgm2 (and may be equal to 4.6x10-9 m/kgm2).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the fan-gearbox axial distance divided by the moment of inertia of the fan may be in the range from 5.3x10-10 m/kgm2 to 2.8x10-9 m/kgm2, and optionally in the range from 8.0x10-10 m/kgm2 to 1.8x10-9 m/kgm2 (and may be equal to 1.3x10-9 m/kgm2).
-
- In various embodiments, the first gearbox support strength ratio may be greater than or equal to 7.0x10-3, and optionally greater than or equal to 1.0x10-2 or greater than or equal to 2.0x10-2.
- In various embodiments, the first gearbox support strength ratio may be in the range from 7.0x10-3 to 2.5x10-1, and optionally in the range from 1.0x10-2 to 1.0x10-1, in the range from 7.0x10-3 to 2.0x10-2 or in the range from 2.0x10-2 to 2.5x10-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm or in the range from 330 to 380 cm, the first gearbox support strength ratio may be in the range from 1.9x10-2 to 2.9x10-2 (and may be equal to 2.4x10-2).
-
- In various embodiments, the second gearbox support strength ratio may be greater than or equal to 1.0x10-1, and optionally greater than or equal to 1.5x10-1, greater than or equal to 1.0x10-1 or greater than or equal to 2.5x10-1.
- In various embodiments, the second gearbox support strength ratio may be in the range from 1.0x10-1 to 3.5, and optionally in the range from 1.5x10-1 to 1.7, in the range from 1.0x10-1 to 2.5x10-1 or in the range from 2.5x10-1 to 3.5.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm or in the range from 330 to 380 cm, the second gearbox support strength ratio may be in the range from 2.8x10-1 to 3.8x10-1 (and may be equal to 3.3x10-1).
-
- In various embodiments, the first gearbox support shear stress ratio may be less than or equal to 4.9x101 m-1, and optionally less than or equal to 20 m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the first gearbox support shear stress ratio may be less than or equal to 35 m-1 or less than or equal to 10.0 m-1. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first gearbox support shear stress ratio may be less than or equal to 12 m-1 or less than or equal to 6.0 m-1.
- In various embodiments, the first gearbox support shear stress ratio may be in the range from 0.35 m-1 to 4.9x101 m-1, and optionally in the range from 0.70 m-1 to 20 m-1.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the first gearbox support shear stress ratio may be in the range from 0.70 m-1 to 35 m-1 and optionally in the range from 6.0 m-1 to 10.0 m-1 (and may be equal to 7.8 m-1).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first gearbox support shear stress ratio may be in the range from 0.50 m-1 to 12 m-1, and optionally in the range from 3.0 m-1 to 6.0 m-1 (and may be equal to 5.4 m-1).
-
- In various embodiments, the second gearbox support shear stress ratio may be less than or equal to 4.1x103 rad/m3, and optionally less than or equal to 1.4x103 rad/m3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the second gearbox support shear stress ratio may be less than or equal to 2.9x103 rad/m3 or less than or equal to 6.5x102 rad/m3. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second gearbox support shear stress ratio may be less than or equal to 7.0x102 rad/m3 or less than or equal to 2.5x102 rad/m3.
- In various embodiments, the second gearbox support shear stress ratio may be in the range from 6.6 rad/m3 to 4.1x103 rad/m3, and optionally in the range from 125x101 rad/m3 to 1.4x103 rad/m3.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the second gearbox support shear stress ratio may be in the range from 2.9x101 rad/m3 to 2.9x103 rad/m3 and optionally in the range from 2.5x102 rad/m3 to 6.5x102 rad/m3 (and may be equal to 4.5x102 rad/m3).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second gearbox support shear stress ratio may be in the range from 1.0x101 rad/m3 to 7.0x102 rad/m3, and optionally in the range from 50 rad/m3 to 2.5x102 rad/m3 (and may be equal to 1.5x102 rad/m3).
-
- In various embodiments, the flight cycle ratio may be less than or equal to 3.20, and optionally less than or equal to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85, or optionally less than or equal to 2.75.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the flight cycle ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the flight cycle ratio may be less than or equal to 2.8 or less than or equal to 2.7.
- In various embodiments, the flight cycle ratio may be in the range from 2.10 to 3.20, and optionally in the range from 2.3 to 2.9.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the flight cycle ratio may be in the range from 2.3 to 3.2 and optionally in the range from 2.8 to 3.0 (and may be equal to 2.9).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the flight cycle ratio may be in the range from 2.1 to 2.8, and optionally in the range from 2.3 to 2.7 (and may be equal to 2.6).
-
- In various embodiments, the flight cycle product may be greater than or equal to 1.00x1016 (N/m2)2, and optionally greater than or equal to 2.05x1016 (N/m2)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the flight cycle product may be greater than or equal to 2.0x1016 (N/m2)2. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the flight cycle product may be greater than or equal to 2.2x1016 (N/m2)2.
- In various embodiments, the flight cycle product may be in the range from 1.00x1016 (N/m2)2 to 7.50x1016 (N/m2)2, and optionally in the range from 2.05x1016 (N/m2)2 to 4.9x1016 (N/m2)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the flight cycle product may be in the range from 2.0x1016 (N/m2)2 to 4.9x1016 (N/m2)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the flight cycle product may be in the range from 2.2x1016 (N/m2)2 to 6.2x1016 (N/m2)2.
-
- In various embodiments, the first torque transmission ratio may be less than or equal to 3.2, and optionally less than or equal to 2.95, optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85 or optionally less than or equal to 2.75.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the first torque transmission ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first torque transmission ratio may be less than or equal to 2.8 or less than or equal to 2.7.
- In various embodiments, the first torque transmission ratio may be in the range from 2.1 to 3.2, and optionally in the range from 2.3 to 2.9.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the first torque transmission ratio may be in the range from 2.3 to 3.2 and optionally in the range from 2.8 to 3.0 (and may be equal to 2.9).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first torque transmission ratio may be in the range from 2.1 to 2.8, and optionally in the range from 2.5 to 2.7 (and may be equal to 2.6).
-
- In various embodiments, the first torque transmission product may be greater than or equal to 2.1x109 (Nm)2, and optionally greater than or equal to 3.5x109 (Nm)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the first torque transmission product may be greater than or equal to 4.0x109 (Nm)2. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first torque transmission product may be greater than or equal to 9.0x109 (Nm)2.
- In various embodiments, the first torque transmission product may be in the range from 2.1x109 (Nm)2 to 9.0x1010 (Nm)2, and optionally in the range from 3.5x109 (Nm)2 to 5.2x1010 (Nm)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the first torque transmission product may be in the range from 4.0x109 (Nm)2 to 9.0x109 (Nm)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the first torque transmission product may be in the range from 9.0x109 (Nm)2 to 9.0x1010 (Nm)2.
-
- In various embodiments, the second torque transmission ratio may be less than or equal to 3.2, and optionally less than or equal to 2.95, or optionally less than or equal to 2.9 (or 2.90), optionally less than or equal to 2.85, or optionally less than or equal to 2.75.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the second torque transmission ratio may be less than or equal to 3.2 or less than or equal to 3.0. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second torque transmission ratio may be greater than or equal to 2.8 or less than or equal to 2.7.
- In various embodiments, the second torque transmission ratio may be in the range from 2.1 to 3.2, and optionally in the range from 2.3 to 2.9.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the second torque transmission ratio may be in the range from 2.3 to 3.2 and optionally in the range from 2.8 to 3.0 (and may be equal to 2.9).
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second torque transmission ratio may be in the range from 2.1 to 2.8, and optionally in the range from 2.3 to 2.7 (and may be equal to 2.6).
-
- In various embodiments, the second torque transmission product may be greater than or equal to 4.1x109 (Nm)2, and optionally greater than or equal to 6.1x109 (Nm)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the second torque transmission product may be greater than or equal to 4.1x109 (Nm)2. In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second torque transmission product may be greater than or equal to 1.1x1010 (Nm)2.
- In various embodiments, the second torque transmission product may be in the range from 4.1x109 (Nm)2 to 9.0x1010 (Nm)2, and optionally in the range from 6.1x109 (Nm)2 to 8.0x1010 (Nm)2.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the second torque transmission product may be in the range from 4.1x109 (Nm)2 to 3.9x1010.
- In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the second torque transmission product may be in the range from 1.1x1010 (Nm)2 to 9.0x1010 (Nm)2.
- In the ratios defined above the maximum take-off and cruise conditions are as defined anywhere herein.
-
Figure 19 illustrates anexample aircraft 1000 having agas turbine engine 10 attached to eachwing gas turbine engine 10 is attached via arespective pylon gas turbines 10 may be that of any embodiment described herein. The aircraft shown inFigure 19 is to be understood as the aircraft for which thegas turbine engine 10 of any embodiment or aspect disclosed herein has been designed to be attached. Theaircraft 1000 has a cruise condition corresponding to the cruise conditions defined elsewhere herein and a max take-off condition corresponding to the MTO conditions defined elsewhere herein. - The present disclosure also relates to a
method 2000 of operating a gas turbine engine on an aircraft (e.g. the aircraft ofFigure 19 ). Themethod 2000 is illustrated inFigure 20 . Themethod 2000 comprises operating 2010 thegas turbine engine 10 described elsewhere herein to provide propulsion for the aircraft to which it is mounted under maximum take-off conditions. The method further comprises operating 2020 the gas turbine engine to provide propulsion during cruise conditions. The gas turbine engine is operated such that any of the parameters or ratios defined herein are within the specified ranges. Cruise conditions and max-take off conditions are as defined elsewhere herein. - The torque on the
core shaft 26 may be referred to as the input torque, as this is the torque which is input to thegearbox 30. The torque supplied by theturbine 19 to the core shaft (i.e. the torque on the core shaft) at cruise conditions may be greater than or equal to 10,000 Nm, and optionally greater than or equal to 11,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque on thecore shaft 26 at cruise conditions may be greater than or equal to 10,000 or 11,000 Nm (and optionally may be equal to 12,760 Nm). In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque on thecore shaft 26 at cruise conditions may be greater than or equal to 25,000 Nm, and optionally greater than or equal to 30,000 Nm (and optionally may be equal to 34,000 Nm). - The torque on the core shaft at cruise conditions may be in the range from 10,000 to 50,000 Nm, and optionally from 11,000 to 45,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque on the
core shaft 26 at cruise conditions may be in the range from 10,000 to 15,000 Nm, and optionally from 11,000 to 14,000 Nm (and optionally may be equal to 12,760 Nm). In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque on thecore shaft 26 at cruise conditions may be in the range from 25,000 Nm to 50,000 Nm, and optionally from 30,000 to 40,000 Nm (and optionally may be equal to 34,000 Nm). - Under maximum take-off (MTO) conditions, the torque on the
core shaft 26 may be greater than or equal to 28,000 Nm, and optionally greater than or equal to 30,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque on thecore shaft 26 under MTO conditions may be greater than or equal to 28,000, and optionally greater than or equal to 35,000 Nm (and optionally may be equal to 36,300 Nm). In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque on thecore shaft 26 under MTO conditions may greater than or equal to 70,000 Nm, and optionally greater than or equal to 80,000 or 82,000 Nm (and optionally may be equal to 87,000 Nm). - Under maximum take-off (MTO) conditions, the torque on the
core shaft 26 may be in the range from 28,000 Nm to 135,000 Nm, and optionally in the range from 30,000 to 110,000 Nm. In some embodiments, for example in embodiments in which the fan diameter is in the range from 240 to 280 cm, the torque on thecore shaft 26 under MTO conditions may be in the range from 28,000 to 50,000 Nm, and optionally from 35,000 to 38,000 Nm (and optionally may be equal to 36,300 Nm). In some embodiments, for example in embodiments in which the fan diameter is in the range from 330 to 380 cm, the torque on thecore shaft 26 under MTO conditions may be in the range from 70,000 Nm to 135,000 Nm, and optionally from 80,000 to 90,000 Nm or 82,000 to 92,000 Nm (and optionally may be equal to 87,000 Nm). - Torque has units of [force]x[distance] and may be expressed in units of Newton metres (N.m), and is defined in the usual way as would be understood by the skilled person.
-
Figure 21 illustrates how the stiffnesses defined herein may be measured.Figure 21 shows a plot of the displacement δ resulting from the application of a load L (e.g. a force, moment or torque) applied to a component for which the stiffness is being measured. At levels of load from zero to LR there is a non-linear region in which displacement is caused by motion of the component (or relative motion of separate parts of the component) as it is loaded, rather than deformation of the component; for example moving within clearance between parts. At levels of load above LS the elastic limit of the component has been exceeded and the applied load no longer causes elastic deformation - plastic deformation or failure of the component may occur instead. Between points R and S the applied load and resulting displacement have a linear relationship. The stiffnesses defined herein may be determined by measuring the gradient of the linear region between points R and S (with the stiffness being the inverse of that gradient). The gradient may be found for as large a region of the linear region as possible to increase the accuracy of the measurement by providing a larger displacement to measure. For example, the gradient may be found by applying a load equal to or just greater than LR and equal to or just less than LS. Although the displacement is referred to as δ in this description, the skilled person would appreciate that equivalent principles would apply to a linear or angular displacement. - The stiffnesses defined herein, unless otherwise stated, are for the corresponding component(s) when the engine is off (i.e. at zero speed/on the bench). The stiffness generally does not vary over the operating range of the engine; the stiffness at cruise conditions of the aircraft to which the engine is used (those cruise conditions being as defined elsewhere herein) may therefore be the same as for when the engine is not in use. However, where the stiffness varies over the operating range of the engine, the stiffnesses defined herein are to be understood as being values for when the engine is at room temperature and unmoving. Values for component strength (e.g. the torsional strength of the gearbox support) as given herein are also at room temperature unless otherwise stated.
- Anything described herein with reference to a planetary type gearbox can apply equally to a star type gearbox unless otherwise stated or where it is apparent that a feature is specific to a particular gearbox type.
- It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.
the method comprising operating the gas turbine engine to provide propulsion for the aircraft under cruise conditions.
Claims (15)
- A gas turbine engine (10) for an aircraft comprising:an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor;a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades;a gearbox (30) that receives an input from the core shaft (26) and outputs drive to a fan shaft (36) via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planet gears (32), a ring gear (38), and a planet carrier (34) arranged to have the plurality of planet gears (32) mounted thereon; anda gearbox support (40) arranged to at least partially support the gearbox (30) within the engine,wherein:the moment of inertia of the fan (23) is greater than or equal to 7.40x107 kgm2; andis greater than or equal to 2.5x10-2 Nkg-1m-3.
- The gas turbine engine (10) according to claim 1, wherein the radial bending stiffness to moment of inertia ratio is greater than or equal to 0.05 Nkg-1m-3, in the range from 2.5x10-2 Nkg-1m-3 to 6.0 Nkg-1m-3, in the range from 0.05 Nkg-1m-3 to 3.0 Nkg-1m-3, or in the range from 0.05 Nkg-1m-3 to 0.6 Nkg-1m-3
- The gas turbine engine (10) according to claim 1 or claim 2, wherein:a) a ratio of the radial bending stiffness of the fan shaft at the output of the gearbox to the moment of inertia of the fan is greater than or equal to 2.5x10-2 Nkg-1m-3, greater than or equal to 0.05 Nkg-1m-3, in the range from 2.5x10-2 Nkg-1m-3 to 6.0 Nkg-1m-3, in the range from 0.05 Nkg-1m-3 to 3.0 Nkg-1m-3, or in the range from 0.05 Nkg-1m-3 to 0.6 Nkg-1m-3; and/orb) a ratio of the radial bending stiffness of the gearbox support to the moment of inertia of the fan is greater than or equal to 3.0x10-2 Nkg-1m-3, greater than or equal to 0.06 Nkg-1m-3, in the range from 3.0x10-2 Nkg-1m-3 to 4.0 Nkg-1m-3, in the range from 0.06 Nkg-1m-3 to 2.0 Nkg-1m-3, or in the range from 0.06 Nkg-1m-3 to 0.48 Nkg-1m-3; and/orc) the radial bending stiffness of the fan shaft (26) at the output of the gearbox is greater than or equal to 4.00x106 N/m, greater than or equal to 3.7x107 N/m, in the range from 4.00x106 N/m to 1.50x109 N/m, or in the range from 3.7x107 N/m to 1.0x109 N/m; and/ ord) the radial bending stiffness of the gearbox support (40) is greater than or equal to 1.0x107 N/m, greater than or equal to 2.0x107 N/m, greater than or equal to 3.0x107 N/m, in the range from 1.0x107 N/m to 4.0x108 N/m, in the range from 2.0x107 N/m to 3.0x108 N/m, or in the range from 3.0x107 N/m to 2.0x108 N/m.
- The gas turbine engine (10) according to any preceding claim, wherein:a) the diameter of the fan is in the range from 240 cm to 280 cm, and the radial bending stiffness to moment of inertia ratio is greater than or equal to 0.05 Nkg-1m-3 or in the range from 0.05 Nkg-1m-3 to 4.0 Nkg-1m-3; ora) the diameter of the fan is in the range from 330 cm to 380 cm and the radial bending stiffness to moment of inertia ratio is greater than or equal to 0.025 Nkg-1m-3 or in the range from 0.025 Nkg-1m-3 to 2.0 Nkg-1m-3.
- The gas turbine engine according to any preceding claim, wherein a product of:
- The gas turbine engine (10) according to any preceding claim, wherein a tilt stiffness to moment of inertia ratio of:
- A gas turbine engine (10) for an aircraft comprising:an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor;a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades;a gearbox (30) that receives an input from the core shaft (26) and outputs drive to a fan shaft (36) via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planet gears (32), a ring gear (38), and a planet carrier (34) arranged to have the plurality of planet gears (32) mounted thereon; anda gearbox support (40) arranged to at least partially support the gearbox (30) within the engine,wherein:the moment of inertia of the fan (23) is greater than or equal to 7.40x107 kgm2; andis greater than or equal to 4.0x10-4 Nrad-1kg-1m-1.
- The gas turbine engine (10) according to claim 7, wherein:the tilt stiffness to moment of inertia ratio is greater than or equal to 1.0x10-3 Nrad-1kg-1m-1, in the range from 4.0x10-4 Nrad-1kg-1m-1 to 2.7x10-1 Nrad-1kg-1m-1, in the range from 1.0x10-3 Nrad-1kg-1m-1 to 0.1 Nrad-1kg-1m-1, or in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.5x10-2 Nrad-1kg-1m-1; and/ora ratio of the tilt stiffness of the fan shaft at the output of the gearbox to the moment of inertia of the fan is greater than or equal to 4.0x10-4 Nrad-1kg-1m-1, greater than or equal to 1.0x10-1 Nrad-1kg-1m-1, in the range from 4.0x10-4 Nrad-1kg-1m-1 to 2.7x10-1 Nrad-1kg-1m-1, in the range from 1.0x10-1 Nrad-1kg-1m-1 to 0.1 Nrad-1kg-1m-1, or in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.5x10-2 Nrad-1kg-1m-1; and/ora ratio of the tilt stiffness of the gearbox support to the moment of inertia of the fan is greater than or equal to 1.0x10-3 Nrad-1kg-1m-1. greater than or equal to 2.0x10-3 Nrad-1kg-1m-1. in the range from 1.0x10-3 Nrad-1kg-1m-1 to 7.0x10-2 Nrad-1kg-1m-1, in the range from 2.0x10-3 Nrad-1kg-1m-1 to 3.0x10-2 Nrad-1kg-1m-1, or in the range from 2.0x10-3 Nrad-1kg-1m-1 to 7.0x10-3 Nrad-1kg-1m-1; and/orthe tilt stiffness of the fan shaft (36) at the output of the gearbox is greater than or equal to 7.00x104 Nm/rad, greater than or equal to 9.5x105 Nm/rad, in the range from 7.00x104 Nm/rad to 7.00x107 Nm/rad, or in the range from 9.5x105 Nm/rad to 3.5x107 Nm/rad; and/orthe tilt stiffness of the gearbox support (40) is greater than or equal to 1.2x105 Nm/rad, greater than or equal to 2.4x105 Nm/rad, greater than or equal to 3.9x105 Nm/rad, in in the range from 1.2x105 Nm/rad to 2.1x107 Nm/rad, in the range from 2.4x105 Nm/rad to 1.6x107 Nm/rad, or in the range from 3.9x105 Nm/rad to 9.0x106 Nm/rad.
- The gas turbine engine (10) according to claim 7 or claim 8, wherein a radial bending stiffness to moment of inertia ratio of:
- The gas turbine engine (10) according to any of claims 7 to 9, wherein:a) the diameter of the fan is in the range from 240 cm to 28 cm, and the tilt stiffness to moment of inertia ratio greater is than or equal to 1.0x10-3 Nrad-1kg-1m-1 or in the range from 1.0x10-3 Nrad-1kg-1m-1 to 1.45x10-2 Nrad-1kg-1m-1; ora) the diameter of the fan is in the range from 330 cm to 380 cm, and the tilt stiffness to moment of inertia ratio is greater than or equal to 4.0x10-4 Nrad-1kg-1m-1 or in the range from 4.0x10-4 Nrad-1kg-1m-1 to 3.0x10-2 Nrad-1kg-1m-1.
- The gas turbine engine according to any of claim 7 to 10, wherein a product of:
- The gas turbine engine (10) according to any preceding claim, wherein:
the moment of inertia of the fan is greater than or equal to 8.3x107 kgm2, in the range from 7.40x107 kgm2 to 9.00x108 kgm2, or in the range from 8.3x107 kgm2 to 6.5x108 kgm2; and/or eithera) the fan blades are formed from a metallic material, optionally titanium or an aluminium-lithium with a titanium leading edge; orb) the fan blades are formed at least partly from an organic matrix composite, optionally having a carbon composite body with a metallic leading edge. - The gas turbine engine according to any preceding claim, wherein:the fan shaft (36) connects the output of the gearbox (30) to the fan (23);a gearbox output position is defined as the point of connection between the fan shaft (36) and the gearbox (30);the fan has an axial centreline;a fan-gearbox axial distance is defined as the axial distance between the axial position of the gearbox output position and the fan axial centreline; andthe fan-gearbox axial distance multiplied by the moment of inertia of the fan is greater than or equal to 1.9x107 kgm3, greater than or equal to 2.9x107 kgm3, in the range from 1.9x107 kgm3 to 6.2x108 kgm3, or in the range from 2.9x107 kgm3 to 3.9x108 kgm3.
- The gas turbine engine (10) according to any preceding claim, wherein:i) the gearbox has a gear ratio of 3.2 to 4.5, and optionally 3.2 to 4.0; and/orii) the gas turbine engine has a specific thrust from 70 to 90 NKg-1; and/oriii) the gas turbine engine has a bypass ratio at cruise conditions of 12.5 to 18, and optionally 13 to 16; and/oriv) the fan (23) has a fan diameter greater than 240 cm and less than or equal to 380 cm, and optionally greater than 300 cm and less than or equal to 380 cm.
- A method (2000) of operating a gas turbine engine (10) for an aircraft, the gas turbine engine comprising:an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor;a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades;a gearbox (30) that receives an input from the core shaft (26) and outputs drive to a fan shaft (36) via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft, the gearbox (30) being an epicyclic gearbox (30) comprising a sun gear (28), a plurality of planet gears (32), a ring gear (38), and a planet carrier (34) arranged to have the plurality of planet gears (32) mounted thereon; anda gearbox support (40) arranged to at least partially support the gearbox (30) within the engine,wherein the moment of inertia of the fan (23) is greater than or equal to 7.40x107 kgm2, and wherein:a) a radial bending stiffness to moment of inertia ratio of:is greater than or equal to 4.0x10-4 Nrad-1kg-1m-1,
the method (2000) comprising operating (2020) the gas turbine engine (10) to provide propulsion for the aircraft under cruise conditions, and wherein:the method optionally further comprises driving the gearbox with an input torque of:i) greater than or equal to 10,000 Nm, and optionally of 10,000 to 50,000 Nm at cruise; and/orii) greater than or equal to 28,000 Nm, and optionally of 28,000 to 135,000 Nm at max-take off conditions.
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
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GBGB1917767.4A GB201917767D0 (en) | 2019-12-05 | 2019-12-05 | High power epicyclic gearbox and operation thereof |
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Publication Number | Publication Date |
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EP3832105A1 true EP3832105A1 (en) | 2021-06-09 |
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EP20206415.0A Withdrawn EP3832105A1 (en) | 2019-12-05 | 2020-11-09 | High power epicyclic gearbox and operation thereof |
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EP (1) | EP3832105A1 (en) |
CN (1) | CN112923026A (en) |
GB (1) | GB201917767D0 (en) |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0924381A2 (en) * | 1997-12-22 | 1999-06-23 | General Electric Company | Frequency tuned turbomachine blade |
US20130336791A1 (en) * | 2011-06-08 | 2013-12-19 | United Technologies Corporation | Geared Architecture for High Speed and Small Volume Fan Drive Turbine |
US20170342858A1 (en) * | 2011-06-08 | 2017-11-30 | United Technologies Corporation | Flexible support structure for a geared architecture gas turbine engine |
US10436035B1 (en) * | 2018-07-03 | 2019-10-08 | Rolls-Royce Plc | Fan design |
US20190316488A1 (en) * | 2018-04-17 | 2019-10-17 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine engine |
-
2019
- 2019-12-05 GB GBGB1917767.4A patent/GB201917767D0/en not_active Ceased
-
2020
- 2020-11-09 EP EP20206415.0A patent/EP3832105A1/en not_active Withdrawn
- 2020-11-10 CN CN202011244291.0A patent/CN112923026A/en active Pending
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0924381A2 (en) * | 1997-12-22 | 1999-06-23 | General Electric Company | Frequency tuned turbomachine blade |
US20130336791A1 (en) * | 2011-06-08 | 2013-12-19 | United Technologies Corporation | Geared Architecture for High Speed and Small Volume Fan Drive Turbine |
US20170342858A1 (en) * | 2011-06-08 | 2017-11-30 | United Technologies Corporation | Flexible support structure for a geared architecture gas turbine engine |
US20190316488A1 (en) * | 2018-04-17 | 2019-10-17 | Rolls-Royce Deutschland Ltd & Co Kg | Gas turbine engine |
US10436035B1 (en) * | 2018-07-03 | 2019-10-08 | Rolls-Royce Plc | Fan design |
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GB201917767D0 (en) | 2020-01-22 |
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